A combustor is a component or area of a gas turbine, ramjet, or scramjet engine where combustion takes place. It is also known as a burner, burner can, combustion chamber or flame holder. In a gas turbine engine, the combustor or combustion chamber is fed high-pressure air by the compression system. The combustor then heats this air at constant pressure as the fuel/air mix burns. As it burns the fuel/air mix heats and rapidly expands. The burned mix is exhausted from the combustor through the nozzle guide vanes to the turbine. In the case of a ramjet or scramjet engines, the exhaust is directly fed out through the nozzle.
A combustor must contain and maintain stable combustion despite very high air flow rates. To do so combustors are carefully designed to first mix and ignite the air and fuel, and then mix in more air to complete the combustion process. Early gas turbine engines used a single chamber known as a can-type combustor. Today three main configurations exist: can, annular, and cannular (also referred to as can-annular tubo-annular). Afterburners are often considered another type of combustor.
Combustors play a crucial role in determining many of an engine's operating characteristics, such as fuel efficiency, levels of emissions, and transient response (the response to changing conditions such as fuel flow and air speed).
The objective of the combustor in a gas turbine is to add energy to the system to power the turbines, and produce a high-velocity gas to exhaust through the nozzle in aircraft applications. As with any engineering challenge, accomplishing this requires balancing many design considerations, such as the following:
Advancements in combustor technology focused on several distinct areas; emissions, operating range, and durability. Early jet engines produced large amounts of smoke, so early combustor advances, in the 1950s, were aimed at reducing the smoke produced by the engine. Once smoke was essentially eliminated, efforts turned in the 1970s to reducing other emissions, like unburned hydrocarbons and carbon monoxide (for more details, see the Emissions section below). The 1970s also saw improvement in combustor durability, as new manufacturing methods improved liner (see Components below) lifetime by nearly 100 times that of early liners. In the 1980s combustors began to improve their efficiency across the whole operating range; combustors tended to be highly efficient (99%+) at full power, but that efficiency dropped off at lower settings. Development over that decade improved efficiencies at lower levels. The 1990s and 2000s saw a renewed focus on reducing emissions, particularly nitrogen oxides. Combustor technology is still being actively researched and advanced, and much modern research focuses on improving the same aspects.[3]
The second type of fuel injector is the air blast injector. This injector "blasts" a sheet of fuel with a stream of air, atomizing the fuel into homogeneous droplets. This type of fuel injector led to the first smokeless combustors. The air used is just some of the primary air (see Air flow paths below) that is diverted through the injector, rather than the swirler. This type of injector also requires lower fuel pressures than the pressure atomizing type.[15]
The vaporizing fuel injector, the third type, is similar to the air blast injector in that primary air is mixed with the fuel as it is injected into the combustion zone. However, the fuel-air mixture travels through a tube within the combustion zone. Heat from the combustion zone is transferred to the fuel-air mixture, vaporizing some of the fuel (mixing it better) before it is combusted. This method allows the fuel to be combusted with less thermal radiation, which helps protect the liner. However, the vaporizer tube may have serious durability problems with low fuel flow within it (the fuel inside of the tube protects the tube from the combustion heat).[16]
The premixing/prevaporizing injectors work by mixing or vaporizing the fuel before it reaches the combustion zone. This method allows the fuel to be very uniformly mixed with the air, reducing emissions from the engine. One disadvantage of this method is that fuel may auto-ignite or otherwise combust before the fuel-air mixture reaches the combustion zone. If this happens the combustor can be seriously damaged.[17]
Can combustors are self-contained cylindrical combustion chambers. Each "can" has its own fuel injector, igniter, liner, and casing.[23] The primary air from the compressor is guided into each individual can, where it is decelerated, mixed with fuel, and then ignited. The secondary air also comes from the compressor, where it is fed outside of the liner (inside of which is where the combustion is taking place). The secondary air is then fed, usually through slits in the liner, into the combustion zone to cool the liner via thin film cooling.[24]
In most applications, multiple cans are arranged around the central axis of the engine, and their shared exhaust is fed to the . Can-type combustors were most widely used in early gas turbine engines, owing to their ease of design and testing (one can test a single can, rather than have to test the whole system). Can-type combustors are easy to maintain, as only a single can needs to be removed, rather than the whole combustion section. Most modern gas turbine engines (particularly for aircraft applications) do not use can combustors, as they often weigh more than alternatives. Additionally, the pressure drop across the can is generally higher than other combustors (on the order of 7%). Most modern engines that use can combustors are turboshafts featuring centrifugal compressors.[25] [26]
The next type of combustor is the "can-annular"[27] combustor. Like the can-type combustor, can-annular combustors have discrete combustion zones contained in separate liners with their own fuel injectors. Unlike the can combustor, all the combustion zones share a common ring (annulus) casing. Each combustion zone no longer has to serve as a pressure vessel.[28] The combustion zones can also "communicate" with each other via liner holes or connecting tubes that allow some air to flow circumferentially. The exit flow from the can-annular combustor generally has a more uniform temperature profile, which is better for the turbine section. It also eliminates the need for each chamber to have its own igniter. Once the fire is lit in one or two cans, it can easily spread to and ignite the others. This type of combustor is also lighter than the can type, and has a lower pressure drop (on the order of 6%). However, a can-annular combustor can be more difficult to maintain than a can combustor.[29] Examples of gas turbine engines utilizing a can-annular combustor include the General Electric J79 turbojet and the Pratt & Whitney JT8D and Rolls-Royce Tay turbofans.[30]
The final, and most-commonly used type of combustor is the fully annular combustor. Annular combustors do away with the separate combustion zones and simply have a continuous liner and casing in a ring (the annulus). There are many advantages to annular combustors, including more uniform combustion, shorter size (therefore lighter), and less surface area.[31] [32] Additionally, annular combustors tend to have very uniform exit temperatures. They also have the lowest pressure drop of the three designs (on the order of 5%).[33] The annular design is also simpler, although testing generally requires a full size test rig. An engine that uses an annular combustor is the CFM International CFM56. Almost all of the modern gas turbine engines use annular combustors; likewise, most combustor research and development focuses on improving this type.
One variation on the standard annular combustor is the double annular combustor (DAC). Like an annular combustor, the DAC is a continuous ring without separate combustion zones around the radius. The difference is that the combustor has two combustion zones around the ring; a pilot zone and a main zone. The pilot zone acts like that of a single annular combustor, and is the only zone operating at low power levels. At high power levels, the main zone is used as well, increasing air and mass flow through the combustor. GE's implementation of this type of combustor focuses on reducing and emissions.[34] A good diagram of a DAC is available from Purdue. Extending the same principles as the double annular combustor, triple annular and "multiple annular" combustors have been proposed and even patented.[35] [36]
One of the driving factors in modern gas turbine design is reducing emissions, and the combustor is the primary contributor to a gas turbine's emissions. Generally speaking, there are five major types of emissions from gas turbine engines: smoke, carbon dioxide (CO2), carbon monoxide (CO), unburned hydrocarbons (UHC), and nitrogen oxides (NOx).[37] [38]
Smoke is primarily mitigated by more evenly mixing the fuel with air. As discussed in the fuel injector section above, modern fuel injectors (such as airblast fuel injectors) evenly atomize the fuel and eliminate local pockets of high fuel concentration. Most modern engines use these types of fuel injectors and are essentially smokeless.[37]
Carbon dioxide is a product of the combustion process, and it is primarily mitigated by reducing fuel usage. On average, 1 kg of jet fuel burned produces 3.2 kg of CO2. Carbon dioxide emissions will continue to drop as manufacturers improve the overall efficiency of gas turbine engines.[38]
Unburned-hydrocarbon (UHC) and carbon-monoxide (CO) emissions are highly related. UHCs are essentially fuel that was not completely combusted. They are mostly produced at low power levels (where the engine is not burning all the fuel).[38] Much of the UHC content reacts and forms CO within the combustor, which is why the two types of emissions are heavily related. As a result of this close relation, a combustor that is well optimized for CO emissions is inherently well optimized for UHC emissions, so most design work focuses on CO emissions.[37]
Carbon monoxide is an intermediate product of combustion, and it is eliminated by oxidation. CO and OH react to form CO2 and H. This process, which consumes the CO, requires a relatively long time ("relatively" is used because the combustion process happens incredibly quickly), high temperatures, and high pressures. This fact means that a low-CO combustor has a long residence time (essentially the amount of time the gases are in the combustion chamber).[37]
Like CO, Nitrogen oxides (NOx) are produced in the combustion zone. However, unlike CO, it is most produced during the conditions that CO is most consumed (high temperature, high pressure, long residence time). This means that, in general, reducing CO emissions results in an increase in NOx, and vice versa. This fact means that most successful emission reductions require the combination of several methods.[37]
See main article: Afterburner. An afterburner (or reheat) is an additional component added to some jet engines, primarily those on military supersonic aircraft. Its purpose is to provide a temporary increase in thrust, both for supersonic flight and for takeoff (as the high wing loading typical of supersonic aircraft designs means that take-off speed is very high). On military aircraft the extra thrust is also useful for combat situations. This is achieved by injecting additional fuel into the jet pipe downstream of (i.e. after) the turbine and combusting it. The advantage of afterburning is significantly increased thrust; the disadvantage is its very high fuel consumption and inefficiency, though this is often regarded as acceptable for the short periods during which it is usually used.
Jet engines are referred to as operating wet when afterburning is being used and dry when the engine is used without afterburning. An engine producing maximum thrust wet is at maximum power or max reheat (this is the maximum power the engine can produce); an engine producing maximum thrust dry is at military power or max dry.
As with the main combustor in a gas turbine, the afterburner has both a case and a liner, serving the same purpose as their main combustor counterparts. One major difference between a main combustor and an afterburner is that the temperature rise is not constrained by a turbine section, therefore afterburners tend to have a much higher temperature rise than main combustors.[39] Another difference is that afterburners are not designed to mix fuel as well as primary combustors, so not all the fuel is burned within the afterburner section.[40] Afterburners also often require the use of flameholders to keep the velocity of the air in the afterburner from blowing the flame out. These are often bluff bodies or "vee-gutters" directly behind the fuel injectors that create localized low-speed flow in the same manner the dome does in the main combustor.[41]
See main article: Ramjet. Ramjet engines differ in many ways from traditional gas turbine engines, but most of the same principles hold. One major difference is the lack of rotating machinery (a turbine) after the combustor. The combustor exhaust is directly fed to a nozzle. This allows ramjet combustors to burn at a higher temperature. Another difference is that many ramjet combustors do not use liners like gas turbine combustors do. Furthermore, some ramjet combustors are dump combustors rather than a more conventional type. Dump combustors inject fuel and rely on recirculation generated by a large change in area in the combustor (rather than swirlers in many gas turbine combustors).[42] That said, many ramjet combustors are also similar to traditional gas turbine combustors, such as the combustor in the ramjet used by the RIM-8 Talos missile, which used a can-type combustor.[43]
See main article: Scramjet.
Scramjet (supersonic combustion ramjet) engines present a much different situation for the combustor than conventional gas turbine engines (scramjets are not gas turbines, as they generally have few or no moving parts). While scramjet combustors may be physically quite different from conventional combustors, they face many of the same design challenges, like fuel mixing and flame holding. However, as its name implies, a scramjet combustor must address these challenges in a supersonic flow environment. For example, for a scramjet flying at Mach 5, the air flow entering the combustor would nominally be Mach 2. One of the major challenges in a scramjet engine is preventing shock waves generated by combustor from traveling upstream into the inlet. If that were to happen, the engine may unstart, resulting in loss of thrust, among other problems. To prevent this, scramjet engines tend to have an isolator section (see image) immediately ahead of the combustion zone.[44]
