• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    It is a system and method for cooling a gas turbine engine [10] which has an axially arranged multistage compressor [52, 54] with paired rotary blades [56, 58] and stationary vanes [60,62 ] located between an external compressor casing [82] and an internal compressor casing [80], and comprising a closed loop compressor cooling by extracting compressor air [92] from a first location on the compressor, cooling it. taking the extracted air [94] and introducing the cooled extracted air [94] into a second location in the compressor, where the second location is upstream of the first location.
    公开号:BR102016026606A2
    申请号:R102016026606-8
    申请日:2016-11-14
    公开日:2018-05-02
    发明作者:El Hacin Sennoun Mohammed；Scott Bunker Ronald
    申请人:General Electric Company；
    IPC主号:
    专利说明:

    (54) Title: METHOD FOR OPERATING A GAS TURBINE ENGINE COMPRESSOR (51) Int. Cl .: F02C 7/00; F04B 53/00 (30) Unionist Priority: 12/03/2015 US 14 / 957,910 (73) Holder (s): GENERAL ELECTRIC COMPANY (72) Inventor (s): MOHAMMED EL HACIN SENNOUN; RONALD SCOTT BUNKER (74) Attorney (s): ANA PAULA SANTOS CELIDONIO (57) Summary: This is a system and method for cooling a gas turbine engine [10] that has an axially arranged multi-stage compressor [ 52, 54] of paired rotating blades [56, 58] and stationary vanes [60.62] located between an external compressor housing [82] and an internal compressor housing [80], and comprising a closed circuit cooling of the compressor, extracting compressor air [92] from a first location on the compressor, cooling the extracted air [94] and introducing the cooled extracted air [94] to a second location on the compressor, where the second location is upstream from the first location.
    1/9 "METHOD FOR OPERATING A GAS TURBINE ENGINE COMPRESSOR"
    Background of the Invention [001] Turbine engines, and particularly gas turbine or combustion engines, are rotary engines that extract energy from a flow of gases through the engine in a series of compressor stages, which includes pairs from rotating blades and stationary vanes, to a combustor and then to a multitude of turbine stages. In the compressor stages, the blades are supported by posts that protrude from the rotor while the blades are mounted in a stator housing. Gas turbine engines have been used for land and nautical locomotion and for power generation, but are most commonly used for aeronautical applications, such as airplanes, including helicopters. In aircraft, gas turbine engines are used to propel the aircraft.
    [002] Gas turbine engines for aircraft are designed to run at high temperatures to maximize engine thrust and therefore cooling certain engine components, such as vanes, is necessary during operation. It is desirable to increase the thermal capacity of the compressor to perform the desired thermal regulation of the engine system.
    Brief Description of the Invention [003] In one aspect, the embodiments of the invention relate to a method for cooling a gas turbine engine compressor that has multiple axially arranged stages of paired rotating blades and stationary vanes located between a compressor housing external and an internal compressor housing. The method includes extracting compressor air from a stage downstream of the compressor, then passing the extracted air through a heat exchanger to cool the extracted air and then introducing the extracted air
    Petition 870160067206, of 11/14/2016, p. 44/60
    2/9 cooled in a stage upstream of the compressor.
    [004] In another aspect, the realizations of the invention refer to a compressor for a gas turbine engine that includes multiple stages axially arranged in flow of paired rotating blades and stationary vanes, a heat exchanger and a cooling circuit that passes by the heat exchanger. An inlet, fluidly coupled to a compressor stage, and an outlet, fluidly coupled to a stage upstream from the stage, allow the compressor air to be extracted through the inlet, passed through the heat exchanger for cooling, after the cooled extracted air is then introduced through the outlet to the upstream stage.
    Brief Description of the Figures [005] In the Figures:
    Figure 1 is a schematic sectional view of a gas turbine engine according to an embodiment of the invention;
    - Figure 2 is a schematic of a compression section of the gas turbine engine of Figure 1 with inter-cooling of some of the compressor stages;
    - Figure 3 is another embodiment of Figure 2;
    - Figure 4 is an additional embodiment of Figure 2; and
    - Figure 5 is a flow chart of a method for cooling a gas turbine engine.
    Description of the Realizations of the Invention [006] The described realizations of the present invention are directed to systems, methods and other devices related to airflow routing in a turbine engine. For purposes of illustration, the present invention will be described in relation to an aircraft gas turbine engine. However, it will be understood that the invention is not limited to this and may have general applicability in non-aeronautical applications, such as
    Petition 870160067206, of 11/14/2016, p. 45/60
    3/9 other mobile applications and industrial, commercial and residential non-mobile applications.
    [007] Figure 1 is a schematic cross-sectional diagram of a gas turbine engine, which can comprise a high-displacement 10 gas turbine engine, for an aircraft. The motor 10 has a geometrical axis which generally extends longitudinally or a central line 12 which extends from the front 14 to the rear 16. The motor 10 includes, in relation to the downstream series flow, a section fan 18 that includes a fan 20, a compressor section 22 that includes an intensifier or low pressure compressor (LP) 24 and a high pressure compressor (HP) 26, a combustion section 28 that includes a combustor 30, a turbine section 32 which includes an HP 34 turbine and an LP 36 turbine and an exhaust section 38. The compressor section 22, the combustion section 28 and the turbine section 32 are in axial flow arranged and wrapped within a core housing 46. The compressor is not limited to an axial orientation and can be oriented axially, radially or in a combined manner.
    [008] The fan section 18 includes a fan housing 40 that surrounds the fan 20. The fan 20 includes a plurality of fan blades 42 arranged radially around the center line 12. The HP compressor 26, the combustor 30 and the HP 34 turbine forms a core 44 of the engine 10 that generates combustion gases. The core 44 is surrounded by the core housing 46 which can be coupled to the fan housing 40. At least a portion of the fan housing 40 surrounds the core housing 46 to define an annular bypass channel 47.
    [009] An HP 48 shaft or coil arranged (or disposed) coaxially around the center line 12 of the motor 10 in an actionable way connects the HP 34 turbine to the HP 26 compressor. An axis or coil of the LP 50, which is arranged (or arranged) coaxially around the central line
    Petition 870160067206, of 11/14/2016, p. 46/60
    4/9 of the motor 10 inside the coil of HP annular larger diameter 48, connectable in an actionable way the turbine of LP 36 to the compressor of LP 24 and to the fan 20. The portions of the motor 10 assembled and that rotate with one or both coils 48, 50 are also referred to individually or collectively as a rotor 51.
    [010] The LP compressor 24 and the HP compressor 26 include, respectively, a plurality of compressor stages 52, 54, in which a set of compressor blades 56, 58 rotates relative to a corresponding set of compressor blades static 60, 62 (also called nozzles) to compress or pressurize the fluid stream that passes through the stage. In a single compressor stage 52, 54, multiple compressor blades 56, 58 can be provided in a ring and can extend radially outwardly relative to the center line 12, from a blade platform to a blade tip, while the corresponding static compressor vanes 60, 62 are positioned downstream and adjacent to the rotating blades 56, 58. Note that the number of blades, vanes and compressor stages shown in Figure 1 were selected for illustration only and that other numbers are possible. Blades 56, 58 for a compressor stage can be mounted on a disc 53, which is mounted on one of the corresponding HP and LP coils 48, 50, with each stage having its own disc. The vanes 60, 62 are mounted on the core housing 46 in a circumferential arrangement around the rotor 51.
    [011] The HP 34 turbine and the LP 36 turbine include, respectively, a plurality of turbine stages 64, 66, wherein a set of turbine blades 68, 70 is rotated relative to a set of turbine blades corresponding static 72, 74 (also called a nozzle) to extract energy from the fluid stream that passes through the stage. In a single turbine stage 64, 66, multiple turbine blades 68, 70 can be
    Petition 870160067206, of 11/14/2016, p. 47/60
    5/9 provided in a ring and can extend radially outwardly from the center line 12, from a paddle platform to a paddle tip, while the corresponding static turbine vanes 72, 74 are positioned upstream and adjacent to rotating blades 68, 70. Note that the number of blades, vanes and turbine stages shown in Figure 1 were selected only by way of illustration and that other numbers are possible.
    [012] The LP compressor 24 and the HP compressor 26 can additionally include at least one guide vane which can be an input vane 55 positioned at the upstream end of the compressor section 22 and an output guide vane 57 positioned at the downstream end of the compressor section 22. The vanes are not limited to one type and can be, for example, non-variable stator vanes or stator vanes.
    [013] In operation, the rotating fan 20 supplies ambient air to the LP compressor 24, which then supplies pressurized ambient air to the HP 26 compressor, which additionally pressurizes the ambient air. The pressurized air in the HP 26 compressor is mixed with fuel in combustion 30 and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP 34 turbine, which drives the HP 26 compressor. The flue gases are discharged into the LP 36 turbine, which extracts additional work to drive the LP 24 compressor and the exhaust gas is finally, discharged from engine 10 through exhaust section 38. The drive of the LP 36 turbine drives the LP 50 coil to turn the fan 20 and the LP compressor 24.
    [014] A portion of the ambient air supplied by the fan 20 can bypass the motor core 44 as a bypass air flow and be used to cool the portions, especially the hot portions, of the motor 10 and / or
    Petition 870160067206, of 11/14/2016, p. 48/60
    6/9 used to cool or power other aspects of the aircraft. In the context of a turbine engine, the hot portions of the engine are usually downstream from the combustor 30, especially the turbine section 32, where the HP 34 turbine is the warmest portion, as it is directly downstream of the combustion section 28.
    [015] Hot engine portions also exist inside the compressor section 22 and, therefore, the ambient air supplied by the fan 20 or cooler air from the compressor can be used, however, without limitation to cooling portions of the compressor section 22. The bypass air flow can pass through a heat exchanger 76, which can be any suitable device, for example, a surface cooler, which has a flat shape or a brick-shaped cooler, which comprises a bar and heat exchanger. plate heat or a tube and shell cooler, which has a cylindrical shape, which can be located within the bypass air flow of bypass channel 47. Although illustrated within bypass channel 47, the location of the heat exchanger 76 is not limited to the bypass channel and can be located in any suitable position within the engine 10, for example, adjacent to an input or output straw 55, 57.
    [016] Referring to Figure 2, a diagram of the compressor section 22 further illustrates an internal compressor housing 80 comprising rotor 51, and an external compressor housing 82 disposed within core housing 46. The multiple stages axially 52, 54 of paired rotating blades 56, 58 and stationary vanes 60, 62 are located between the external compressor housing 82 and the internal compressor housing 80. A cooling circuit 78 comprises an inlet 84 fluidly coupled to a first compressor portion 26. Inlet 84 is located adjacent to external compressor housing 82, adjacent to blade tip 58, and adjacent to end wall 86
    Petition 870160067206, of 11/14/2016, p. 49/60
    7/9 of the vane 62. The first portion can be in a stage downstream 54 of the compressor 26 or in any suitable area upstream of the blades 58 for the downstream stage 54 and downstream of the vane 62 for the upstream stage 52.
    [017] Cooling circuit 78 further comprises an outlet 88 fluidly coupled to a second portion located upstream of the first portion of compressor 26. Output 88 is located adjacent to external compressor housing 82, adjacent to the blade tip 56 , and adjacent to the end wall 86 of the vane 60. The second portion may be in an upstream stage 52 of the compressor 24 or in any suitable area upstream of the blades 56 for the upstream stage 52 and upstream of the vanes 62 for the downstream stage 52.
    [018] The first and second portions can be in any relative downstream / upstream location along the compressor, at any stage in the compressor as long as the relative position is correct. The first and second portions can be at different stages of the compressor or at the same stage of the compressor. For example, as shown in Figure 3, the first portion and the second portion may be portions adjacent to the vanes and blades within the same compressor stage 54. The first and second portions may also be coupled to an interior of any of the elements airfoil, such as reeds. For example, as shown in Figure 4, the first portion is downstream of a blade 58 and the second portion is upstream of the blade 58 and coupled to the inside 90 of a reed 60.
    [019] Figure 5 further illustrates a method 200 for cooling a gas turbine engine 10 using the cooling circuit 78 that passes through heat exchanger 76. In a first step 202, compressor air 92 is extracted from the first portion of compressor 26 through the inlet
    Petition 870160067206, of 11/14/2016, p. 50/60
    8/9
    84. Secondarily 204 the compressor air 92 flows through the heat exchanger 76 to form the cooled extracted air 94. In a final step 206, the cooled extracted air 94 is introduced through outlet 88 to the second portion located upstream of the first portion of the compressor 26.
    [020] The heat exchanger 76 cools the compressor air 92 to form the cooled extracted air 94, which in turn is used to cool areas in the upstream stage 52 of the compressor 26 which include, however, are not limited to interior 90 of vanes 60, 62. In yet another embodiment, heat exchanger 76 can be used to transfer energy to or from a closed loop cooling system in which the closed loop coolant is then used for other thermal adjustments for heating or cooling purposes.
    [021] Figures 2 to 4 include a cooling fluid that can be fan air 95. Fan air 95 passes through heat exchanger 76 to cool compressor air 92. Extracted compressor air 92 can also be air alternatively routed 96 directed to the fan or guide vane region and passed through other heat exchangers, such as brick coolers and surface coolers.
    [022] An optional flow control device, for example, but without limitation is a control valve that can be included in the handle, so that the coolant flow can be either on, off or modulated depending on conditions of operation.
    [023] This written description uses examples to reveal the invention, including the best way, and also to allow anyone skilled in the art to practice the invention, including making and using any devices or systems and performing any built-in methods. The patentable scope of the invention is defined by the claims and may include other examples that occur to those versed
    Petition 870160067206, of 11/14/2016, p. 51/60
    9/9 in technique. Such other examples are intended to be within the scope of the claims, if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.
    Petition 870160067206, of 11/14/2016, p. 52/60
    1/2
    权利要求:
    Claims (10)
    [1]
    Claims
    1. METHOD FOR OPERATING A GAS TURBINE ENGINE [24, 26] COMPRESSOR [10] that has multiple stages [52, 54] of paired rotating blades [56, 58] and stationary vanes [60, 62] located between one external compressor housing [82] and an internal compressor housing [80], in which the method is characterized by the fact that it comprises:
    - extract compressor air [92] from a stage downstream [52, 54] from the compressor [24, 26];
    - pass the extracted air through a heat exchanger to cool the extracted air; and
    - introduce the cooled extracted air into an upstream stage [52, 54] of the compressor [24, 26].
    [2]
    2. METHOD, according to claim 1, characterized by the fact that the extraction comprises extracting compressor air [92] upstream of the blades [56, 58] for the downstream stage [52, 54] and downstream of the vanes [60, 62] for the upstream internship [52, 54].
    [3]
    3. METHOD, according to claim 1, characterized by the fact that the extraction comprises extracting compressor air [92] adjacent to the external compressor housing [82].
    [4]
    4. METHOD, according to claim 3, characterized by the fact that the extraction comprises extracting the air adjacent to a tip of the blades [56, 58] of the downstream stage [52, 54].
    [5]
    5. METHOD, according to claim 3, characterized by the fact that the extraction comprises extracting the air adjacent to a vane end wall [60, 62] of the downstream stage [52, 54].
    [6]
    6. METHOD, according to claim 1, characterized by the fact that the introduction comprises introducing the cooled extracted air [94]
    Petition 870160067206, of 11/14/2016, p. 53/60
    2/2 upstream of the blades [56, 58] for the upstream stage [52, 54] and upstream of the vanes [60, 62] for the downstream stage [52, 54].
    [7]
    7. METHOD, according to claim 1, characterized by the fact that the introduction comprises introducing the cooled extracted air [94] adjacent to the external compressor housing [82].
    [8]
    8. METHOD, according to claim 7, characterized by the fact that the introduction comprises introducing the cooled extracted air [94] adjacent to a blade tip [56, 58] of the upstream stage [52, 54].
    [9]
    9. METHOD, according to claim 7, characterized by the fact that the introduction comprises introducing the cooled extracted air [94] adjacent to a vane end wall [60, 62] of the upstream stage [52, 54].
    [10]
    10. METHOD, according to claim 1, characterized by the fact that the introduction comprises introducing the cooled extracted air [94] through an interior of the vanes [60, 62] of the upstream stage [52, 54].
    Petition 870160067206, of 11/14/2016, p. 54/60
    1/5 't
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    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US3220697A|1963-08-30|1965-11-30|Gen Electric|Hollow turbine or compressor vane|
    US5611197A|1995-10-23|1997-03-18|General Electric Company|Closed-circuit air cooled turbine|
    DE102007026455A1|2007-06-05|2008-12-11|Rolls-Royce Deutschland Ltd & Co Kg|Jet engine with compressor air circulation and method of operating the same|
    US8858161B1|2007-11-29|2014-10-14|Florida Turbine Technologies, Inc.|Multiple staged compressor with last stage airfoil cooling|
    DE102013211042B3|2013-06-13|2014-08-14|MTU Aero Engines AG|Compressor operating method for aero engine gas turbine, involves controlling air valve arrangement based on detected operating mode, and cooling air flow by coolant valve arrangement through injecting arrangement|
    WO2015122949A2|2013-12-17|2015-08-20|United Technologies Corporation|Adaptive turbomachine cooling system|US10731560B2|2015-02-12|2020-08-04|Raytheon Technologies Corporation|Intercooled cooling air|
    US10371055B2|2015-02-12|2019-08-06|United Technologies Corporation|Intercooled cooling air using cooling compressor as starter|
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    US10480419B2|2015-04-24|2019-11-19|United Technologies Corporation|Intercooled cooling air with plural heat exchangers|
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    US10794288B2|2015-07-07|2020-10-06|Raytheon Technologies Corporation|Cooled cooling air system for a turbofan engine|
    US10330010B2|2015-12-14|2019-06-25|United Technologies Corporation|Compressor core inner diameter cooling|
    US10443508B2|2015-12-14|2019-10-15|United Technologies Corporation|Intercooled cooling air with auxiliary compressor control|
    US10669940B2|2016-09-19|2020-06-02|Raytheon Technologies Corporation|Gas turbine engine with intercooled cooling air and turbine drive|
    US10794290B2|2016-11-08|2020-10-06|Raytheon Technologies Corporation|Intercooled cooled cooling integrated air cycle machine|
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    US10577964B2|2017-03-31|2020-03-03|United Technologies Corporation|Cooled cooling air for blade air seal through outer chamber|
    US10711640B2|2017-04-11|2020-07-14|Raytheon Technologies Corporation|Cooled cooling air to blade outer air seal passing through a static vane|
    US10738703B2|2018-03-22|2020-08-11|Raytheon Technologies Corporation|Intercooled cooling air with combined features|
    US10808619B2|2018-04-19|2020-10-20|Raytheon Technologies Corporation|Intercooled cooling air with advanced cooling system|
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    US10718233B2|2018-06-19|2020-07-21|Raytheon Technologies Corporation|Intercooled cooling air with low temperature bearing compartment air|
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    DE102019208731A1|2019-06-14|2020-12-17|Rolls-Royce Deutschland Ltd & Co Kg|Engine with bleed air return for the compressor and method for operating an engine|
    法律状态:
    2018-05-02| B03A| Publication of a patent application or of a certificate of addition of invention [chapter 3.1 patent gazette]|
    2019-11-05| B08F| Application dismissed because of non-payment of annual fees [chapter 8.6 patent gazette]|Free format text: REFERENTE A 3A ANUIDADE. |
    2020-02-27| B08K| Patent lapsed as no evidence of payment of the annual fee has been furnished to inpi [chapter 8.11 patent gazette]|Free format text: EM VIRTUDE DO ARQUIVAMENTO PUBLICADO NA RPI 2548 DE 05-11-2019 E CONSIDERANDO AUSENCIA DE MANIFESTACAO DENTRO DOS PRAZOS LEGAIS, INFORMO QUE CABE SER MANTIDO O ARQUIVAMENTO DO PEDIDO DE PATENTE, CONFORME O DISPOSTO NO ARTIGO 12, DA RESOLUCAO 113/2013. |
    优先权:
    申请号 | 申请日 | 专利标题
    US14/957,910|US20170159568A1|2015-12-03|2015-12-03|Intercooling system and method for a gas turbine engine|
    US14/957,910|2015-12-03|
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