Abstract:
A gas turbine engine has a compressor section received within an inner housing. An is an outer housing is spaced radially outwardly of the inner core housing. A nacelle has an anti-icing system which taps compressed air from the compressor section through an anti-ice valve and to the nacelle. The anti-ice valve is opened at startup of the gas turbine engine to assist compressor stability.

Description:
BACKGROUND OF THE INVENTION 
     This application relates to a gas turbine engine wherein a nacelle anti-ice valve provides a startup bleed valve function. 
     Gas turbine engines are known, and typically include a fan delivering air into a bypass duct defined within a nacelle, and also into a core engine. The air in the core engine flow passes through a compressor section, and then into a combustion section. In the combustion section the air is mixed with fuel and ignited, and products of this combustion pass downstream over turbine rotors. 
     There are many challenges in the design of a gas turbine engine. One challenge occurs at startup. There is typically a large load on the compressor as it begins rotating at startup. Thus, it is known to have a bleed valve in place that opens to allow the partially compressed air to be dumped out of the compressor section. In many engines, there are a plurality of these bleed valves. 
     It is also known to provide a nacelle anti-icing system. The nacelle anti-icing system typically will tap hot air from the compressor section, and selectively deliver it to the lip of the nacelle to provide anti-icing at the lip of the nacelle. This anti-icing function is performed selectively, and is not necessary during much of the operation of a gas turbine engine on an aircraft. However, when conditions indicate that there may be icing at the lip of the nacelle, the valve may be opened to deliver the hot air to that location. 
     In the prior art, the use of plural compressor stability bleed valves increases the complexity of the system. Further, should one of these bleed valves fail, air would be continuously bled from the compressor section. This would be undesirable, as the efficiency of the engine would be reduced and the hot air could damage other components positioned in the core. 
     SUMMARY OF THE INVENTION 
     In a featured embodiment, a gas turbine engine has a compressor section received within an inner housing. An outer housing is spaced radially outwardly of the inner housing. A nacelle is also included. A nacelle anti-icing system taps compressed air from the compressor section through an anti-ice valve and to the nacelle. The anti-ice valve is opened at startup of the gas turbine engine for the purpose of compressor stability assistance. 
     In another embodiment according to the previous embodiment, the anti-ice valve is normally open, but may be closed by a control. 
     In another embodiment according to any of the previous embodiments, the anti-ice system includes a nozzle positioned adjacent an upstream lip of the nacelle. 
     In another embodiment according to any of the previous embodiments, a compressor stability bleed valve is positioned in the inner housing for selectively dumping air that has been at least partially compressed. The bleed valve is also opened at startup. 
     In another embodiment according to any of the previous embodiments, the anti-ice valve is opened at startup of the gas turbine engine, without regard to ambient conditions. 
     In another embodiment according to any of the previous embodiments, a fan is included in the gas turbine engine, and delivers air into a bypass duct inwardly of the nacelle, and also into the compressor section. 
     In another embodiment according to any of the previous embodiments, a bypass ratio can be described as the volume of air passing into the bypass duct compared to the volume of air passing into the compressor. The bypass ratio is greater than about 6. 
     In another embodiment according to any of the previous embodiments, the bypass ratio is greater than about 10. 
     In another embodiment according to any of the previous embodiments, the fan is driven by a turbine that is included in the gas turbine engine. A gear reduction is positioned between the fan and turbine. 
     In another embodiment according to any of the previous embodiments, a gear ratio of the gear reduction is greater than about 2.3. 
     In another embodiment according to any of the previous embodiments, the gear reduction is greater than about 2.5. 
     In another embodiment according to any of the previous embodiments, a fan is included in the gas turbine engine, and delivers air into a bypass duct inwardly of the nacelle, and also into the compressor section. 
     In another embodiment according to any of the previous embodiments, a bypass ratio can be described as the volume of air passing into the bypass duct compared to the volume of air passing into the compressor. The bypass ratio is greater than about 6. 
     In another embodiment according to any of the previous embodiments, the bypass ratio is greater than about 10. 
     In another embodiment according to any of the previous embodiments, the fan is driven by a turbine that is included in the gas turbine engine. A gear reduction is positioned between the fan and turbine. 
     In another embodiment according to any of the previous embodiments, a gear ratio of the gear reduction is greater than about 2.3. 
     In another embodiment according to any of the previous embodiments, the gear reduction is greater than about 2.5. 
     In another featured embodiment, a gas turbine engine has a compressor section received within an inner housing. An outer housing is spaced radially outwardly of the inner housing. A nacelle is also included. A compressor stability bleed valve in the inner housing selectively dumps air that has been at least partially compressed in the compressor section into a space between the inner and outer housings. A nacelle anti-icing system taps compressed air from the compressor section through an anti-ice valve and to the nacelle. The anti-ice valve and bleed valve are opened at startup of the gas turbine engine. The anti-ice valve is normally open, but may be closed by a control. The anti-ice system includes a nozzle positioned adjacent an upstream lip of the nacelle. The anti-ice valve is opened at startup without regard to ambient conditions. 
     In another embodiment according to the previous embodiment, a fan is included in the gas turbine engine, and delivers air into a bypass duct inwardly of the nacelle, and also into the compressor section. 
     In another embodiment according to any of the previous embodiments, a bypass ratio can be described as the volume of air passing into the bypass duct compared to the volume of air passing into the compressor. The bypass ratio is greater than about 6. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  schematically shows a gas turbine engine. 
         FIG. 2  is a cross-section through a high pressure compressor section. 
         FIG. 3  shows details of a gas turbine engine. 
     
    
    
     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 in a bypass duct inwardly of a nacelle  80 . The compressor section  24  receives 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 pressure 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 , directly or through a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The high pressure 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 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 [(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. 
       FIG. 2  shows the compressor section  52  having an inner wall  81 , and an outer wall  82 . These features may be part of a gas turbine engine generally as disclosed in  FIG. 1 . 
     The compressor section is provided with a compressor stability bleed valve  94 . This valve is under the control of a control  196  which selectively opens the bleed valve  94  during engine startup such that compressed air is dumped outwardly of the compressor section  52  into a chamber  96 . This reduces the load on the compressor rotors as they begin to rotate. As can be appreciated, the compressor stability bleed valve  94  dumps air into the chamber  96 , and thus, components  200 , shown schematically, within the space  96  are exposed to this hot air. 
     The control  196  also controls an anti-ice valve  88 . The anti-ice valve  88  also includes a tap  86  for tapping compressed air. As would be understood by someone who works in this art, this compressed air would be hot. 
     As shown in  FIG. 3 , the tap  86  passes through the anti-ice valve  88 , into a conduit  84 , and then to a nozzle  90  associated with a lip  92  at an upstream end of the nacelle  80 . 
     The nozzle  90  would shoot air in opposed circumferential directions such that the lip  92  is provided with this hot air, should conditions indicate that there may be icing. Typically, the anti-ice valve  88  would not be left open at all times, as that would reduce the efficiency of the compressor. 
     In the prior art, the anti-ice valve  88  is normally closed, however, a control will open the valve when conditions indicate icing. In general, the anti-ice valve  88  has remained closed at startup, when the compressor stability bleed valves might open. In some cases, an anti-ice valve may have been opened at startup, but only if ambient conditions dictated the use. The present control algorithm would ensure the anti-ice valve is opened at startup, without consideration of ambient conditions. Further, while the specific embodiment does include both a bleed valve  94 , and the anti-ice valve  88 , it is possible the anti-ice valve  88  could be utilized on its own within the scope of this disclosure. 
     In the present application, the control  196  may open the anti-ice valve  88  at startup. Alternatively, the anti-ice valve  88  may be designed such that it is normally opened, and is left open at startup. In such an arrangement, the control  196  would be operable to close the valve  88  when conditions do not warrant the tapping of hot air for an anti-icing function. That is, the anti-ice valve is opened in an unactuated state, but can be actuated to be closed. 
     Thus, the present invention utilizes the anti-ice valve  88  to perform not only the anti-ice function, but also to provide a compressor stability bleed valve. This thus eliminates the need for plural bleed valves. Further, should the valve  88  fail, it is directing hot air to a less sensitive area than does bleed valve  94 . 
     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.