US20130192195A1 - Gas turbine engine with compressor inlet guide vane positioned for starting - Google Patents

Gas turbine engine with compressor inlet guide vane positioned for starting Download PDF

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Publication number
US20130192195A1
US20130192195A1 US13/367,742 US201213367742A US2013192195A1 US 20130192195 A1 US20130192195 A1 US 20130192195A1 US 201213367742 A US201213367742 A US 201213367742A US 2013192195 A1 US2013192195 A1 US 2013192195A1
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US
United States
Prior art keywords
compressor
gas turbine
turbine engine
vane
set forth
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Abandoned
Application number
US13/367,742
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English (en)
Inventor
Eric J. Wehmeier
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Raytheon Technologies Corp
Original Assignee
United Technologies Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Technologies Corp filed Critical United Technologies Corp
Priority to US13/367,742 priority Critical patent/US20130192195A1/en
Assigned to UNITED TECHNOLOGIES CORPORATION reassignment UNITED TECHNOLOGIES CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: WEHMEIER, ERIC J.
Priority to PCT/US2013/021799 priority patent/WO2013154638A1/fr
Priority to SG11201402934YA priority patent/SG11201402934YA/en
Priority to EP13775616.9A priority patent/EP2809921B1/fr
Publication of US20130192195A1 publication Critical patent/US20130192195A1/en
Priority to US14/259,180 priority patent/US11208950B2/en
Abandoned legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/06—Varying effective area of jet pipe or nozzle
    • F02K1/12—Varying effective area of jet pipe or nozzle by means of pivoted flaps
    • F02K1/1207—Varying effective area of jet pipe or nozzle by means of pivoted flaps of one series of flaps hinged at their upstream ends on a fixed structure
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • F02C3/13—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor having variable working fluid interconnections between turbines or compressors or stages of different rotors
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D19/00—Starting of machines or engines; Regulating, controlling, or safety means in connection therewith
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/26—Starting; Ignition
    • F02C7/262—Restarting after flame-out
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F02C9/16—Control of working fluid flow
    • F02C9/20—Control of working fluid flow by throttling; by adjusting vanes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/075—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type controlling flow ratio between flows
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/85—Starting
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2270/00—Control
    • F05D2270/01—Purpose of the control system
    • F05D2270/09—Purpose of the control system to cope with emergencies
    • F05D2270/092—Purpose of the control system to cope with emergencies in particular blow-out and relight

Definitions

  • This application relates to a gas turbine engine having an inlet guide vane which has its position controlled to increase windmilling speed of engine components.
  • Gas turbine engines typically include a fan delivering air into a bypass duct outwardly of a core engine, and into a compressor in the core engine. Air in the compressor is passed downstream into a combustor section where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors, driving them, and in turn drive the compressor and fan. Recently it has been proposed to include a gear reduction between a low pressure compressor and the fan, such a low pressure turbine can drive the two at distinct speeds.
  • a gas turbine engine as used on an aircraft must be able to start under several conditions. First, the gas turbine engine must be able to start when on the ground. A starter can be used on the ground. Second, the gas turbine engine must be able to start in the air. In the air, at lower speeds of the aircraft, the normal starter for the gas turbine engine may be utilized to begin driving the turbine/compressor rotors. However, at higher speeds the starter may not be utilized. At higher speeds so called “windmilling” is relied upon at startup. Windmilling typically occurs as the compressor and fan rotors are driven by the air being forced into the core engine, and the bypass duct, as the aircraft continues to move.
  • a gas turbine engine has a compressor section, a low spool, and a fan.
  • the fan delivers air into the compressor section.
  • the compressor section compresses air and delivers it into a combustion section.
  • the combustion section mixes air with fuel, ignites the fuel, and drives the products of the combustion across turbine rotors.
  • the compressor section includes a variable inlet guide vane which is movable between distinct angles to control the airflow approaching the compressor section.
  • a control for the gas turbine engine is programmed to position the vane at startup of the engine to increase airflow across the compressor section.
  • the compressor section includes a low pressure compressor and a high pressure compressor.
  • the vane is positioned forwardly of an upstream most rotor in the low pressure compressor.
  • the fan is driven with the low pressure compressor by the low spool. There is a gear reduction between the fan and the low spool.
  • control includes stored desired positions for the vane to provide increased airflow into the compressor at startup at various conditions.
  • the various conditions include the altitude of an aircraft carrying the gas turbine engine, and an air speed of the aircraft.
  • the conditions also include a speed of a low spool which rotates with the low pressure compressor when startup is occurring.
  • the fan also delivers bypass air into a bypass duct position outwardly of a core engine including the compressor, the combustor and the turbine rotors.
  • the bypass duct has a variable area nozzle, and the position of the nozzle also is controlled at startup to increase airflow through the bypass duct and across the fan.
  • a bypass ratio of the volume of air passes into the bypass duct to the volume delivered into the compressor section is greater than about 6.
  • the fan also delivers bypass air into a bypass duct position outwardly of a core engine including the compressor, the combustor and the turbine rotors.
  • the bypass duct has a variable area nozzle. The position of the nozzle also is controlled at startup to increase airflow through the bypass duct and across the fan.
  • a bypass ratio of the volume of air passes into the bypass duct to the volume delivered into the compressor section is greater than about 6.
  • the position of the vane is selected to increase airflow across the low pressure compressor while an aircraft associated with the gas turbine engine is in the air, and to increase windmilling speed of the low and high spools.
  • a starter is also utilized in combination with the windmilling while the aircraft is in the air to start the engine.
  • a method of starting a gas turbine engine includes the steps of providing a variable inlet guide vane forwardly of a compressor, and moving the vane to a position at startup of the engine selected to increase airflow across the compressor, and starting the gas turbine engine.
  • the desired position is selected to provide increased airflow into the compressor at startup at various conditions.
  • the various conditions include the altitude of an aircraft carrying the gas turbine engine, and the airspeed of the aircraft.
  • the compressor includes a low pressure compressor and a high pressure compressor.
  • the conditions also include a speed of a spool which drives the low pressure compressor.
  • a fan delivers bypass air into a bypass duct positioned outwardly of a core engine which includes the compressor, the bypass duct having a variable area nozzle, and the position of the nozzle being controlled at startup to increase airflow across the fan.
  • a variable inlet guide vane has an inlet guide vane provided with an actuator to change an angle of the vane.
  • a control for the actuator is programmed to position the vane at a position to increase airflow across the vane at startup of an associated aircraft
  • FIG. 1 shows a gas turbine engine
  • FIG. 2 is a schematic of a control logic circuit.
  • FIG. 3 is a flowchart.
  • FIG. 1 schematically illustrates a gas turbine engine 20 .
  • the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22 , a compressor section 24 , a combustor section 26 and a turbine section 28 .
  • Alternative engines might include an augmentor section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flowpath B while the compressor section 24 drives air along a core flowpath C for compression and communication into the combustor section 26 then expansion through the turbine section 28 .
  • FIG. 1 schematically illustrates a gas turbine engine 20 .
  • the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22 , a compressor section 24 , a combustor section 26 and a turbine section 28 .
  • Alternative engines might include an augmentor section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flowpath B while the compressor section 24 drives air along
  • the 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.
  • 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 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 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54 .
  • a mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46 .
  • the mid-turbine frame 57 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 C is compressed by the low pressure compressor 44 then the high pressure compressor 52 , mixed and burned with fuel in the combustor 56 , then expanded over the high pressure turbine 54 and low pressure turbine 46 .
  • the mid-turbine frame 57 includes airfoils 59 which are in the core airflow path.
  • the turbines 46 , 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
  • the engine 20 in one example is a high-bypass geared aircraft engine.
  • the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than 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
  • the low pressure turbine 46 has a pressure ratio that is greater than about 5.
  • 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
  • the low pressure turbine 46 has a pressure ratio that is greater than about 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 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
  • 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.
  • TSFC Thrust Specific Fuel Consumption
  • 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 [(Tambient deg R)/518.7) ⁇ 0.5].
  • the “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second.
  • the gas turbine engine 20 is provided with controls and features to optimize starting.
  • a starter 400 is typically included with a gas turbine engine, and is relied upon to begin driving the low spool and high spool when the engine is started. This will typically occur when the airplane is on the ground, and is a relatively simple process at that time.
  • the starter may be utilized while the aircraft is in the air to begin driving rotation of the low and high spool 32 to begin the restart process.
  • the combustion section has begun to ignite and burn the fuel, then the products of combustion will take over driving the turbine rotors and the starter may stop.
  • the engine is provided with equipment that is controlled to optimize to increase the ability to maximize windmilling and high spool.
  • an actuator 180 selectively drives a control to position a compressor inlet guide vane 184 which is just forward of the forward most low compressor rotor 186 .
  • An angle of the vane 184 is preferably positioned to maximize the flow of air reaching the rotor 186 while the aircraft is being restarted. In flight, this would be positioning the vane 184 such that the air being forced into the core engine as the aircraft continues to move through the air with engine 20 not being powered, is maximized.
  • bypass airflow B may be maximized by positioning a variable fan nozzle 200 .
  • the variable fan nozzle 200 is controlled by an actuator 204 , shown schematically, to move axially and control the flow area at 202 . Generally, one would open the nozzle to a full open position to maximize this air flow.
  • Both the inlet guide vane 180 and the actuator 204 for the variable area fan nozzle 200 are generally as known. However, they have not been utilized at startup to maximize the amount of windmilling which occurs.
  • Applicant has developed a control system as shown in FIG. 2 which takes in altitude signals 210 , an aircraft speed signal 212 , and a signal 214 which is the windmilling speed of the low spool 30 .
  • Lookup tables are stored in control component 216 , 218 and 222 . Applicant has developed tables which associate particular altitudes, engine speed, and Mach number, with a desired position for the vane 184 , and/or the position of the nozzle 200 to maximize the airflow as discussed above.
  • the desired positions can be developed experimentally and will vary by aircraft and engine design. While the two features may be used in combination, it is also within the scope of this application that each could be used individually without the other, where appropriate.
  • the signal passes downstream to a block 224 , wherein additional second signal comes from control elements 218 and 216 .
  • Elements 216 and 218 provides an adjustment to the output of element 222 based upon the low spool 30 speed altitude and aircraft airspeed.
  • a signal passes to the actuators 180 and/or 204 .
  • the FIG. 2 control can be incorporated into a FADEC 199 .
  • the altitude would be generally the same, and the Mach number would be zero. Further, the low spool speed might be zero. Even so, there would be desired positions for the vane 184 and/or nozzle 200 .
  • a starter 400 shown schematically, in combination with the windmilling However, this would all be incorporated into the lookup tables stored in components 216 , 218 and 222 . Also, as mentioned above, at times the starter 400 cannot be relied upon in some circumstances. Again, this would be anticipated and relied upon at components 216 , 218 and 222 or in the look-up table.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Control Of Turbines (AREA)
US13/367,742 2012-01-31 2012-02-07 Gas turbine engine with compressor inlet guide vane positioned for starting Abandoned US20130192195A1 (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
US13/367,742 US20130192195A1 (en) 2012-01-31 2012-02-07 Gas turbine engine with compressor inlet guide vane positioned for starting
PCT/US2013/021799 WO2013154638A1 (fr) 2012-01-31 2013-01-17 Turbine à gaz dotée d'une aube guide d'entrée de compresseur positionnée pour le démarrage
SG11201402934YA SG11201402934YA (en) 2012-01-31 2013-01-17 Gas turbine engine with compressor inlet guide vane positioned for starting
EP13775616.9A EP2809921B1 (fr) 2012-01-31 2013-01-17 Turbine à gaz dotée d'une aube guide d'entrée de compresseur positionnée pour le démarrage
US14/259,180 US11208950B2 (en) 2012-01-31 2014-04-23 Gas turbine engine with compressor inlet guide vane positioned for starting

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US201261592667P 2012-01-31 2012-01-31
US13/367,742 US20130192195A1 (en) 2012-01-31 2012-02-07 Gas turbine engine with compressor inlet guide vane positioned for starting

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US13/367,742 Abandoned US20130192195A1 (en) 2012-01-31 2012-02-07 Gas turbine engine with compressor inlet guide vane positioned for starting
US14/259,180 Active 2034-10-08 US11208950B2 (en) 2012-01-31 2014-04-23 Gas turbine engine with compressor inlet guide vane positioned for starting

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EP (1) EP2809921B1 (fr)
SG (1) SG11201402934YA (fr)
WO (1) WO2013154638A1 (fr)

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US11668253B2 (en) * 2020-10-16 2023-06-06 Pratt & Whitney Canada Corp. System and method for providing in-flight reverse thrust for an aircraft
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WO2013154638A1 (fr) 2013-10-17
EP2809921A4 (fr) 2015-10-21
EP2809921B1 (fr) 2018-06-13
US20140223916A1 (en) 2014-08-14
US11208950B2 (en) 2021-12-28
EP2809921A1 (fr) 2014-12-10

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