Abstract:
An apparatus for controlling the temperature of a component, which is situated in use in a gas stream, provides a nozzle to create a jet of air at an angle to the gas stream, the jet being directed into the region of the stagnation point of the component so as to control the temperature of the component. The invention is particularly suited to preventing or reducing the formation of ice on vanes of gas turbine engines, but may also be applied to other components, and may equally be used in situations where a component is to be cooled rather than heated.

Description:
CROSS REFERENCE TO RELATED APPLICATION 
     This application is entitled to the benefit of British Patent Application No. GB 0708459.3 filed on May 2, 2007. 
     FIELD OF THE INVENTION 
     This invention relates to components that are situated in use in a gas stream, and whose temperature needs to be controlled. 
     BACKGROUND OF THE INVENTION 
     For example, it is known that certain regions of a gas turbine engine—particularly those nearest the air intake—may be susceptible to ice formation on components such as guide vanes, struts and duct walls. Icing may occur at any time, whether or not the engine is running, if the atmospheric conditions are appropriate. When the engine is running, icing may occur during ground running, at idle or at higher engine speeds, as well as during operation in flight. In such circumstances ice may build up on, and then be shed from, these components, and the ice may cause damage to other components further downstream in the engine. The risk of icing is exacerbated when the design of the engine is such that the fan or low-pressure compressor imparts only a small temperature rise to the air. If ice has built up on components of the engine while it is not running, it may or may not be shed from the engine immediately on starting. 
     In order to avoid ice build-up on vulnerable components it is known to make these components hollow, so that hot air from the combustor or elsewhere in the engine can be used to warm the component and thereby prevent icing. However, hollow components increase engine complexity, and consequently manufacturing costs and timescales. In cases where components have complicated 3D geometry for aerodynamic reasons, as is increasingly common, it may be difficult or impossible to make them hollow. Components, which move or rotate (such as variable vanes), add yet more complexity and potential leakage, because the hot air flow must be provided through a rotating spindle to the component. 
     Furthermore, this method of heating the component relies principally on heat soak through the component walls. It is therefore relatively inefficient, and has the further disadvantage that it tends to heat the whole component, not only that part of it susceptible to icing. 
     It would therefore be desirable to have an improved method of preventing icing of components, which overcomes the disadvantages of known techniques. 
     SUMMARY OF THE INVENTION 
     According to the invention, there is provided an apparatus for controlling the temperature of a component, the component situated in use in a duct in which flows a gas stream, whereby a stagnation point is created in the region of an upstream portion of the component, the apparatus including a nozzle to create a jet of air at an angle to the gas stream, the jet being directed into the region of the stagnation point of the component so as to control the temperature of the component, characterised in that the nozzle is located in a wall of the duct. 
     According to another aspect of the present invention, a gas turbine engine having, in axial flow series, an air intake, a propulsive fan, an intermediate pressure compressor, a high pressure compressor, a combustor, a turbine including a high pressure turbine, an intermediate pressure turbine and a low pressure turbine, and an exhaust nozzle, said turbines including a plurality of components, with each component situated in use in a duct in which flows a gas stream, whereby a stagnation point is created in the region of an upstream portion of the component, the apparatus including a nozzle to create a jet of air at an angle to the gas stream, the jet being directed into the region of the stagnation point of the component so as to control the temperature of the component, characterised in that the nozzle is located in a wall of the duct. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic partial section of a gas turbine engine of known type; 
         FIG. 2  is a schematic sectional view of an apparatus according to the invention; 
         FIG. 3  is an enlarged view on the line III-III in  FIG. 2 ; 
         FIG. 4  is a schematic sectional view of an alternative apparatus according to the invention; and 
         FIG. 5  is an enlarged view on the line V-V in  FIG. 4 . 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     Referring to  FIG. 1 , a gas turbine engine is generally indicated at  10  and comprises, in axial flow series, an air intake  11 , a propulsive fan  12 , an intermediate pressure compressor  13 , a high pressure compressor  14 , a combustor  15 , a turbine comprising a high pressure turbine  16 , an intermediate pressure turbine  17  and a low pressure turbine  18 , and an exhaust nozzle  19 . 
     The gas turbine engine  10  operates in a conventional manner. Air entering the intake  11  is accelerated by the fan  12 , which produces two air flows—a first air flow passes through an annular duct  24  into the intermediate pressure compressor  13  and a second air flow passes through an annular bypass duct  26  and provides propulsive thrust. The intermediate pressure compressor  13  compresses the air flow directed into it before delivering that air to the high pressure compressor  14  where further compression takes place. 
     The compressed air exhausted from the high pressure compressor  14  is directed into the combustor  15  where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive, the high, intermediate and low pressure turbines  16 ,  17  and  18  before being exhausted through the nozzle  19  to provide additional propulsive thrust. The high, intermediate and low pressure turbines  16 ,  17  and  18  respectively drive the high and intermediate pressure compressors  14  and  13  and the fan  12  by concentric shafts  20 ,  21  and  22 , which rotate about an axis X-X. 
       FIG. 2  shows part of the duct  24  indicated in  FIG. 1 . The duct is defined by radially outer and radially inner walls, respectively  26  and  28 , and contains an annular array of vanes  30 , of which one is shown. The vanes  30  are secured between the duct walls  26 ,  28 . 
     In use, a gas stream flows through the duct  24  in the direction shown by the arrow  32 . Under appropriate atmospheric conditions, entrained water or ice crystals in the gas stream may form ice on the leading edge region  34  of the vane  30 . 
     A short distance upstream of the leading edge  34  a nozzle  36  is provided in the duct wall  26 . Air  38  is fed through this nozzle, creating a jet of air  40  directed along the leading edge  34  of the vane  30  and substantially perpendicular to the gas stream direction  32 . Each vane  30  in the annular array is provided with a nozzle  36 . The air feed to the nozzles  36  may be by any convenient means—for example, an annular manifold may surround the duct wall  26  and provide a common air supply for all the nozzles  36 . The air  38  will normally be taken from a cabin or handling bleed offtake on the engine, and so this air will be at a higher temperature than the air in the gas stream  32 . 
       FIG. 3  is a view on the line III-III in  FIG. 2 , looking towards the outer duct wall  26 . The gas stream  32  flows towards the leading edge  34  of the vane  30 . Dashed arrows  42  indicate how the gas stream  32  is deflected by the vane  30 . There is a roughly triangular region  44 , immediately upstream of the leading edge  34 , in which there is relatively little movement of the gas. The nozzle  36  is positioned in this stagnation region  44 . Consequently, the jet of air  40  is not significantly deflected by the gas stream  32 ,  42  and will flow along the leading edge  34  of the vane  30 , warming it and preventing the build-up of ice. (This may be seen more clearly in  FIG. 2 .) 
     Because the jet of air  40  is not significantly deflected by the gas stream  32 ,  42 , the injected air remains close to the boundary layer. Therefore, the heat transfer between the air and the component surface is more efficient, and the temperature increase needed in the component can be achieved with a smaller volume of injected air. If the jet of air  40  were injected elsewhere, so that it mixed with the gas stream  32 ,  42 , a much greater volume of injected air would be needed to achieve the required heating. 
     This apparatus has the advantage over known de-icing methods that only the part of the vane most susceptible to icing is heated, and so the efficiency penalty of providing this heating is minimized. 
       FIGS. 4 and 5  show an alternative embodiment of the invention. An annular duct  124  is bounded by outer and inner walls  126  and  128 , and contains an annular array of variable stator vanes  130  of known type, of which one is shown. The vane  130  may pivot about an axis Z-Z. At its outer end, the vane  130  is mounted on a circular mounting member  146 , commonly known as a penny. The surface of the penny is shaped to conform to the profile of the outer duct wall  126 , and in effect forms a part of that wall. 
     In use, a gas stream flows through the duct  124  in the direction shown by the arrow  132 . The dashed arrows  142  indicate how the gas stream  132  is deflected by the vane  130 . As in the first embodiment, shown in  FIGS. 2 and 3 , there is a stagnation region  144 ; but in contrast to the first embodiment, this region is not adjacent to the leading edge of the vane  130 . A nozzle  136  in the penny  146  is located in the stagnation region  144 , and directs air along the surface of the vane  130  to warm the vane and prevent the build-up of ice. 
     In the two embodiments of the invention that have been described, a jet of air is directed at a vane of a gas turbine engine in order to heat it and prevent the build-up of ice. It will be appreciated, by those skilled in the art, that other embodiments are possible employing the same inventive principle as in these embodiments. 
     In particular, the jet of air may be employed to cool a component rather than to heat it. Referring, for example, to  FIGS. 2 and 3 , if the air  38  were cooler than the gas stream  32  then its effect would be to cool the leading edge region  34  of the vane  30 . The mechanism and advantages of such an apparatus would be exactly the same as those set out above in respect of heating a component, mutatis mutandis. 
     The direction of the gas stream  32 ,  132 , relative to the aerofoil axis, need not be as shown in either of the two embodiments described, but will be dictated by the aerodynamic characteristics of a particular embodiment. 
     It will be appreciated that the exact position of the stagnation region  44 ,  144  relative to the leading edge of the vane  30 ,  130  will then depend on the geometry of the apparatus and the direction of the gas stream  32 ,  132 . 
     Those skilled in the art will recognise that the hole  36 ,  136  may be of any suitable shape, as dictated by the requirements of a particular embodiment. 
     The invention need not be applied only to a vane or a variable vane of a gas turbine engine. It could equally well be used for a strut, or for any component (in a gas turbine, a steam turbine or any other machine) that is located in a gas stream in use and requires heating or cooling. 
     The invention may be used during starting of a gas turbine engine, to melt or shed ice that has built up on components while the engine has not been running. 
     In the first embodiment described, the air for the jet of air  40  is supplied from a bleed offtake of the gas turbine engine. In other embodiments, the air may be supplied from any convenient source within or outside the machine incorporating the component. In particular, where the invention is to be used during the starting of an aircraft gas turbine engine, the air may be supplied by another engine on the aircraft, by an APU or from a ground cart or other external source. Where the invention is used in a steam turbine, the air may be supplied from the low pressure steam circuit or from an external source of compressed air. 
     It may be desirable, in certain applications, to provide more than one nozzle associated with a component. The nozzles may direct the flow of air at different angles, or at different regions of the component, to optimize the heating or cooling effect.