GAS TURBINE ENGINE WITH COMPRESSOR INLET GUIDE VANE POSITIONED FOR STARTING

Abstract

A gas turbine engine includes a variable inlet guide vane positioned forwardly of a low pressure compressor. The angle of the inlet guide vane is controlled at startup to increase airflow into the compressor. This is particularly useful when the gas turbine engine is being restarted while an associated aircraft is in the air, and is relied upon to increase windmill speed of the compressor and turbine rotors. A method and variable inlet vane are also disclosed.

Description

CROSS-REFERENCE TO RELATED APPLICATION

This application claims priority to U.S. Provisional Application No. 61/592,667, which was filed Jan. 31, 2012.

BACKGROUND

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 are known, and 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.

SUMMARY

In a featured embodiment, 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.

In a further embodiment according to the foregoing embodiment, the compressor section includes a low pressure compressor and a high pressure compressor.

In a further embodiment according to the foregoing embodiment, the vane is positioned forwardly of an upstream most rotor in the low pressure compressor.

In a further embodiment according to the foregoing embodiment, 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.

In a further embodiment according to the foregoing embodiment, the control includes stored desired positions for the vane to provide increased airflow into the compressor at startup at various conditions.

In a further embodiment according to the foregoing embodiment, the various conditions include the altitude of an aircraft carrying the gas turbine engine, and an air speed of the aircraft.

In a further embodiment according to the foregoing embodiment, the conditions also include a speed of a low spool which rotates with the low pressure compressor when startup is occurring.

In a further embodiment according to the foregoing embodiment, 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.

In a further embodiment according to the foregoing embodiment, 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.

In a further embodiment according to the foregoing embodiment, 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.

In a further embodiment according to the foregoing embodiment, 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.

In a further embodiment according to the foregoing embodiment, 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.

In a further embodiment according to the foregoing embodiment, a starter is also utilized in combination with the windmilling while the aircraft is in the air to start the engine.

In another featured embodiment, 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.

In a further embodiment according to the foregoing embodiment, the desired position is selected to provide increased airflow into the compressor at startup at various conditions.

In a further embodiment according to the foregoing embodiment, the various conditions include the altitude of an aircraft carrying the gas turbine engine, and the airspeed of the aircraft.

In a further embodiment according to the foregoing embodiment, the compressor includes a low pressure compressor and a high pressure compressor.

In a further embodiment according to the foregoing embodiment, the conditions also include a speed of a spool which drives the low pressure compressor.

In a further embodiment according to the foregoing embodiment, 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.

In another featured embodiment, 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

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a gas turbine engine.

FIG. 2 is a schematic of a control logic circuit.

FIG. 3 is a flowchart.

DETAILED DESCRIPTION

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. Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

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. In a further example, 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 and the low pressure turbine 46 has a pressure ratio that is greater than about 5. 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 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.

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. The flight condition of 0.8 Mach and 35,000 ft, 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 [(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.

On the other hand, there are times when the gas turbine engine is shut down while an aircraft associated with the gas turbine engine is still in the air. At lower air speeds, 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. Of course, once 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.

Under certain conditions, use of the starter while the aircraft is in the air is not advised or is not possible. Under those conditions, the force of air being driven into the engine core, and across the fan 42 is relied upon to drive the turbine rotors, and the compressor rotors. This process is called “windmilling.”

It is desirable to increase the speed of windmilling of the high spool that occurs when it is necessary to restart the engine because higher windmill speeds driver higher airflow.

The engine is provided with equipment that is controlled to optimize to increase the ability to maximize windmilling and high spool. Thus, 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.

Also, the 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.

In general, it is desirable to position the vane 184 to maximize airflow through the core engine, and position the variable area nozzle 200 to maximize airflow across the fan 42. Airflow across the fan 42 will drive the fan to rotate, and air being forced into the core engine will cause the compressor rotor 186 to rotate.

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 control of the area fan nozzle is disclosed in co-pending application entitled Gas Turbine Engine With Variable Area Fan Nozzle Positioned for Starting, filed on even date herewith, Ser. No ______.

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.

Downstream of the block 229, a signal passes to the actuators 180 and/or 204. The FIG. 2 control can be incorporated into a FADEC 199.

Of course, if the aircraft is positioned on the ground, 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. If the aircraft is in the air when being restarted and moving at a relatively slow Mach number, it may be possible to utilize 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.

Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.

Claims

1. A gas turbine engine comprising:

a compressor section;

a low spool;

a fan;

said fan delivering air into said compressor section, and said compressor section compressing air and delivering it into a combustion section;

the combustion section mixing air with fuel, igniting the fuel, and driving the products of the combustion across turbine rotors;

said compressor section including a variable inlet guide vane which is movable between distinct angles to control the airflow approaching said compressor section;

a control for said gas turbine engine, programmed to position said vane at startup of the engine to direct airflow across said compressor section; and

said fan also delivering bypass air into a bypass duct positioned outwardly of a core engine including said compressor section, said combustor, and said turbine rotors, and wherein the position of said vane is configured to direct airflow across the compressor section while an aircraft associated with the gas turbine engine is in the air, and to increase a windmilling speed of the compressor section and said turbine rotors.

2. The gas turbine engine as set forth in claim 1, wherein said compressor section including a low pressure compressor and a high pressure compressor.

3. The gas turbine engine as set forth in claim 2, wherein said vane is positioned forwardly of an upstream most rotor in the low pressure compressor.

4. The gas turbine engine as set forth in claim 3, wherein said fan is driven with said low pressure compressor by said low spool, and there being a gear reduction between said fan and said low spool.

5. The gas turbine engine as set forth in claim 2, wherein said control includes stored desired positions for said vane to provide increased airflow into the compressor at startup at various conditions.

6. The gas turbine engine as set forth in claim 5, wherein said various conditions include the altitude of an aircraft carrying the gas turbine engine, and an air speed of the aircraft.

7. The gas turbine engine as set forth in claim 6, wherein the conditions also include a speed of a low spool, said low spool rotating with said low pressure compressor when startup is occurring.

8. The gas turbine engine as set forth in claim 5, wherein said bypass duct having a variable area nozzle, and the position of the nozzle also being controlled at startup to provide airflow through said bypass duct and across said fan to increase the windmilling speed.

9. The gas turbine engine as set forth in claim 8, wherein a bypass ratio of the volume of air passing into said bypass duct to the volume delivered into said compressor section is greater than about 6.

10. The gas turbine engine as set forth in claim 1, wherein said bypass duct having a variable area nozzle, and the position of the nozzle also being controlled at startup to provide airflow through said bypass duct and across said fan to increase the windmilling speed.

11. The gas turbine engine as set forth in claim 10, wherein a bypass ratio of the volume of air passing into said bypass duct to the volume delivered into said compressor section is greater than about 6.

12. (canceled)

13. The gas turbine engine as set forth in claim 1, wherein a starter is also utilized in combination with the windmilling while the aircraft is in the air to start the engine.

14. A method of starting a gas turbine engine comprising the steps of:

(a) providing a variable inlet guide vane forwardly of a compressor, and moving said vane to a position at startup of the engine selected to provide airflow across said compressor;

(b) starting said gas turbine engine; and

wherein a fan delivers bypass air into a bypass duct positioned outwardly of a core engine which includes said compressor, and turbine rotors, and wherein the position of the vane is configured to direct airflow across the compressor while an aircraft associated with the gas turbine engine is in the air, and to increase a windmilling speed of the compressor and turbine rotors.

15. The method as set forth in claim 14, wherein said desired position is selected to provide increased airflow into the compressor at startup at various conditions.

16. The method as set forth in claim 15, wherein said various conditions include the altitude of an aircraft carrying the gas turbine engine, and the airspeed of the aircraft.

17. The method as set forth in claim 14, wherein said compressor includes a low pressure compressor and a high pressure compressor.

18. The method as set forth in claim 16, wherein the conditions also include a speed of a spool which drives the low pressure compressor.

19. The method as set forth in claim 18, wherein the bypass duct having a variable area nozzle, and the position of the nozzle being controlled at startup to provide airflow across the fan to increase the windmilling speed.

20. A variable inlet guide vane comprising:

an inlet guide vane provided with an actuator to change an angle of the vane; and

a control for the actuator programmed to position said vane at a position to provide airflow across said vane at startup of an associated aircraft, and wherein the vane position is configured to direct airflow across said vane to a compressor, to increase a windmilling speed of the compressor and an associated turbine rotor while an aircraft that receives the variable inlet guide vane is in the air.

21. The gas turbine engine as set forth in claim 1, wherein said vane is positioned at startup to maximize airflow across said compressor section.

22. The method as set forth in claim 14, wherein said vane is positioned to maximize airflow to said compressor.

23. The variable inlet guide vane as set forth in claim 20, wherein said vane is positioned at startup to maximize airflow to the compressor.