Patent Application
20190107059
The subject matter disclosed herein generally relates to combustors in gas turbine engines and, more particularly, to cooling of combustors of gas turbine engines.
A combustor of a gas turbine engine may be configured and required to burn fuel in a minimum volume. Such configurations may place substantial heat load on the structure of the combustor (e.g., panels, shell, etc.). Such heat loads may dictate that special consideration is given to structures which may be configured as heat shields or panels configured to protect the walls of the combustor. Even with such configurations, excess temperatures at various locations may occur leading to oxidation, cracking, and high thermal stresses of the heat shields or panels.
According to one embodiment, a combustor cooling system for a combustor of a gas turbine engine is provided. The combustor cooling system comprising: a heat exchanger fluidly connected to a compressor section of the gas turbine engine, the heat exchanger being configured to receive compressor bleed air from the compressor section and cool the compressor bleed air; and a plenum assembly fluidly connected to the heat exchanger, the plenum assembly configured to receive cooled compressor bleed air from the heat exchanger and distribute the cooled compressor bleed air to one or more areas of the combustor.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the plenum assembly is fluidly connected to the heat exchanger through a valve configured to adjust the flow of cooled compressor bleed air from the heat exchanger to the combustor.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the plenum assembly further comprises at least one of a frontal plenum to provide cooled compressor bleed air to an inlet of the combustor, an inner plenum to provide cooled compressor bleed air to a radially inward area of the combustor, an outer plenum to provide cooled compressor bleed air to a radially outward area of the combustor, and a bypass plenum to provide cooled compressor air to an engine component located downstream of the combustor.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the plenum assembly further comprises a frontal plenum to provide cooled compressor bleed air to an inlet of the combustor, an inner plenum to provide cooled compressor bleed air to a radially inward area of the combustor, an outer plenum to provide cooled compressor bleed air to a radially outward area of the combustor, and a bypass plenum to provide cooled compressor air to an engine component located downstream of the combustor.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the plenum assembly further comprises a frontal plenum to provide cooled compressor bleed air to an inlet of the combustor.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the plenum assembly further comprises an inner plenum to provide cooled compressor bleed air to a radially inward area of the combustor and an outer plenum to provide cooled compressor bleed air to a radially outward area of the combustor.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the plenum assembly further an inner plenum to provide cooled compressor bleed air to a radially inward area of the combustor.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the plenum assembly further comprises an outer plenum to provide cooled compressor bleed air to a radially outward area of the combustor.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the plenum assembly further comprises a bypass plenum to provide cooled compressor air to an engine component downstream of the combustor.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the plenum assembly further an inner plenum to provide cooled compressor bleed air to a radially inward area of the combustor.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the plenum assembly further comprises an outer plenum to provide cooled compressor bleed air to a radially outward area of the combustor.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the plenum assembly is fluidly connected to one or more engine components located downstream of the combustor and the plenum assembly is configured to provide the cooled compressor bleed air to one or more engine components.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the plenum assembly further comprises an inner plenum to provide cooled compressor bleed air to a radially inward area of the combustor, and wherein the inner plenum assembly provides cooled compressor bleed air to an inner vane of a turbine section of the gas turbine engine.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the plenum assembly further comprises an outer plenum to provide cooled compressor bleed air to a radially outward area of the combustor, and wherein the outer plenum assembly provides cooled compressor bleed air to an outer vane of a turbine section of the gas turbine engine or a blade outer air seal of the turbine section of the gas turbine engine.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the plenum assembly further comprises a bypass plenum to provide cooled compressor bleed air to a blade of a turbine section of the gas turbine engine.
According to another embodiment, method of cooling a combustor of a gas turbine engine is provided. The method comprising: extracting compressor bleed air from a compressor section of the gas turbine engine; cooling compressor bleed air using a heat exchanger fluidly connected to the compressor section; and distributing cooled compressor bleed air to one or more areas of the combustor using a plenum assembly.
In addition to one or more of the features described above, or as an alternative, further embodiments may include adjusting flow of cooled compressor bleed air from the heat exchanger to the combustor using a valve.
In addition to one or more of the features described above, or as an alternative, further embodiments may include detecting a temperature of the combustor; determining a cooling requirement for the combustor in response to the temperature of the combustor; and adjusting flow of cooled compressor bleed air from the heat exchanger to the combustor using a valve.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the one or more areas comprises at least one of an inlet of the combustor, a radially inward area of the combustor, and a radially outward area of the combustor.
In addition to one or more of the features described above, or as an alternative, further embodiments may include distributing cooled compressor bleed air to one or more engine components located downstream of the combustor.
The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, that the following description and drawings are intended to be illustrative and explanatory in nature and non-limiting.
The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike:
The detailed description explains embodiments of the present disclosure, together with advantages and features, by way of example with reference to the drawings.
A detailed description of one or more embodiments of the disclosed apparatus and method are presented herein by way of exemplification and not limitation with reference to the Figures.
Combustors of gas turbine engines experience elevated heat levels during operation. Impingement and convective cooling of panels of the combustor wall may be used to help cool the combustor. Convective cooling may be achieved by air that is trapped between the panels and a shell of the combustor. Impingement cooling may be a process of directing relatively cool air from a location exterior to the combustor toward a back or underside of the panels.
Thus, combustor liners and heat shields are utilized to face the hot products of combustion within a combustion chamber and protect the overall combustor shell. The combustor liners may be supplied with cooling air including dilution passages which deliver a high volume of cooling air into a hot flow path. The cooling air may be air from the compressor of the gas turbine engine; however as overall pressure ratio increase to increase efficiency in gas turbine engine, the temperature of the air from the compressor also increases, thus reducing the cooling ability of the air. Embodiments disclosed herein include apparatuses and methods to increase the cooling ability of the air from the compressor.
The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 300 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. An engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The engine static structure 36 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 300, then expanded over the high pressure turbine 54 and low pressure turbine 46. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.
The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 m/sec).
The combustor 300, as shown in
Compressor bleed air 210a may be supplied to the combustor 300 through the combustor cooling system 200. The combustor cooling system 200 comprises a heat exchanger 220, a valve 230, and a plenum assembly 250. The heat exchanger 220 is fluidly connected to the compressor section 24. The heat exchanger 220 is configured to receive the compressor bleed air 210a and cool the compressor bleed air 210a, the heat exchanger 220 will expel cooled compressor bleed air 210b. The heat exchanger 220 is also fluidly connected to the fan section 22 and is configured to receive fan bleed air 440 from the fan section 22. The fan bleed air 440 may provide additional cooling for the heat exchanger 220. It understood that the embodiments disclosed herein are not limited to utilizing fan bleed air 440 to cool the heat exchanger 220 and bleed air from any engine section upstream of the high pressure compressor 52 (i.e. the final compression stage) may be utilized to cool the heat exchanger 220. The valve 230 is configured to regulate airflow from the heat exchanger 220 to the plenum assembly 250 by adjusting the valve 230. The valve 230 may fluidly connect the heat exchanger 220 and the combustor 300 as shown in
The controller 440 may include a processor and memory. For ease of illustration, the processor and memory are not shown in
The plenum assembly 250 is fluidly connected to the heat exchanger 220 and configured to receive cooled compressor bleed air 210b from the heat exchanger 220. The plenum assembly 250 is configured to distribute the cooled compressor bleed air 210b to one or more areas of the combustor 300. The plenum assembly 250 may include at least one of a frontal plenum 260, an inner plenum 270, an outer plenum 280, and a bypass plenum 290. The frontal plenum 260 provides cooled compressor bleed air 210b to the inlet 306 of the combustor 300. The inner plenum 270 provides cooled compressor bleed air 210b to a radially inward area 303 of the combustor 300. The outer plenum 280 provides cooled compressor bleed air 210b to a radially outward area 305 of the combustor 300. A plurality of passages 306 from the plenum assembly 250 to the combustion chamber 302 pass through the panels 326, 328 and shell 300 to enable cooled compressor bleed air 210b to flow into the combustion chamber 302 from the inward plenum 270 and/or the outward plenum 280. The panels 326, 328 and shell 300 may also include a plurality of holes and/or apertures to enable fluid, such as gases, to flow from areas external to the combustion chamber 302 into the combustion chamber 302.
As illustrated, the plenum assembly 250 may also be fluidly connected to one or more engine components 410, 420, 430 located downstream of the combustor 300. The plenum assembly 250 provides cooled compressor bleed air 210b to the one or more engine components 410, 420, 430. The inner plenum 270 may be fluidly connected to a first engine component 410 and provide cooled compressor bleed air 210b to the first engine component 410. The first engine component 410 may be at least one of an inner portion of a vane of the turbine section 28 and a rotor cavity. The outer plenum 280 may be fluidly connected to a second engine component 420 and provide cooled compressor bleed air 210b to the second engine component 420. The second engine component 420 may be an outer portion of a vane of the turbine section 28 or a blade outer air seal. The bypass plenum 290 may be fluidly connected to a third engine component 430 and provide cooled compressor bleed air 210b to the third engine component 430. The third engine component 430 may be at least one of a blade of the turbine section 28, the rotor system, and tangential onboard injector.
The plenum assembly 250 may include various combinations of the frontal plenum 260, the inner plenum 270, the outer plenum 280, and the bypass plenum 290. As shown in
Referring now to
The method 700 may also include distributing cooled compressor bleed air 210b to one or more engine components 410, 420, 430 located downstream of the combustor 300, as discussed above. The method 700 may further include adjusting flow of cooled compressor bleed air 210b from the heat exchanger 220 to the combustor 300 using a valve 220. As described above, a controller 440 may be in electronic communication with the valve 220 and may control the operation of the valve. The controller 440 may detect a temperature of the combustor 300, determine a cooling requirement for the combustor 300 in response to the temperature of the combustor 300; and adjust flow of cooled compressor bleed air 210b from the heat exchanger 220 to the combustor 300 using a valve 220. The controller 440 may detect a temperature of the combustor 300 using either a temperature sensor (not-shown) and/or through calculations.
While the above description has described the flow process of
Technical effects of embodiments of the present disclosure include cooling compressor bleed air with a heat exchanger and utilizing the cooled compressor bleed air to cool a combustor of a gas turbine engine.
The term “about” is intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application. For example, “about” can include a range of ±8% or 5%, or 2% of a given value.
The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof.
While the present disclosure has been described with reference to an exemplary embodiment or embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof. Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.
