Turbomachine rotor blade tip sealing

Abstract: A gas turbine engine turbomachine (10) comprises a turbine section (18) in which cooling air is supplied from a compressor section (14) via a pipe (68) to a chamber (60). The chamber (60) supplies the cooling air through passages (58) in the nozzle guide vanes (30) to impinge upon a stator disc (32). A valve (70) in the pipe (68) is used to switch the flow of cooling air on and off. The valve (70) is opened to allow the cooling air to cool the stator disc (32) to shrink the stator disc (32) to minimise clearance between a shroud (34) supported by the nozzle guide vanes (30) and the turbine blades (28) during one mode of operation. The valve (70)is closed to allow the stator disc (32) to expand to increase the clearance between the shroud (34) and the turbine blades (28).

Description: The present invention relates to turbomachines, particularly to gas turbine engines, more particularly to apparatus for controlling the clearance between rotor blades and a surrounding shroud of gas turbine engine compressors or rotors. In order to increase the efficiency of turbomachines it is well known to maintain the closest possible clearance between the rotor blades and the surrounding shroud to ensure that very little of the working fluid passes through the clearance. It is well known to have increased clearance between the rotor blades and the surrounding shroud to ensure that the rotor blades do not rub on the shroud during transient conditions or during violent manoeuvring of an aircraft to which the gas turbine engine is secured. US Patent No4,849,895 describes a system for adjusting a radial clearance between the tips of the blades of a turbine rotor stage and encircling stator elements such as shroud means. The system controls the radial clearance by means of a valve in the air flow duct activated by a signal based upon real-time calculations carried out by an electronic computer. The clearance control is achieved by varying the amount of cooling air supplied to the turbine section in accordance with the computer calculations. In one arrangement the air is directed at the exterior of the turbine casing in jets from a surrounding air manifold. In another arrangement the air supply from the valve is ducted into an annular chamber around the turbine casing, from which the air passes through passageways in the low pressure turbine and circulates from one stage to another, in known fashion. This known method of allowing cooling air to circulate through the turbine stages is considered inefficient and does not produce the close control of tip clearances required in turbine stages in which the stators are built around a stator disc. The present invention seeks to provide a turbomachine with novel apparatus for controlling the clearance between the rotor blades and a surrounding shroud. According to the present invention there is provided a turbomachine comprising a rotor having a plurality of radially outward extending rotor blades, on one side of the rotor a stator assembly including a stator disc and a plurality of stator vanes arranged in a first stage extending radially outward from the stator disc, shroud means encircling the rotor blades at a radial clearance therefrom, an opposite side of the shroud means being located between the inner and outer radial stops on an adjacent stator assembly to limit radial movement of the shroud means, and a fluid supply means controlled by valve means and having at least one passage extending radially through the stator assembly, characterised in that the shroud means is cantilevered on one side from the first stage of stator vanes, and the at least one passage is arranged to direct a supply of fluid to impinge directly onto the stator disc and the valve means is operated to control the supply of the fluid onto the disc in order to influence the temperature and therefore the diameter of the disc and thereby control the radial clearance between the tips of the rotor blades and the shroud means (34). Preferably the stator disc and the first stage of stator vanes are located downstream of the rotor blades. Preferably the valve means comprises an on/off valve arranged to supply cooling or heating fluid in a first mode of operation and to terminate the supply of cooling or heating fluid in a second mode of operation. Preferably the supply of cooling or heating fluid is a supply of cool fluid to cool the stator disc, the valve means is arranged tc supply cool fluid to the stator disc to cool the stator disc to reduce the clearance between the shroud and the rotor blades during cruise and the valve means is arranged to terminate the supply of cooling fluid to the stator disc to increase the clearance between the shroud and the rotor blades during high speed operation of the turbomachine. Preferably the turbomachine is a gas turbine engine. The rotor may be a turbine rotor and the stator may be a turbine stator. The means to supply cooling or heating fluid may include bleed means to bleed fluid from the compressor of the gas turbine engine. The present invention will be more fully described by way of example with reference to the accompanying drawings in which: Figure 1 is a partially cut away view of a gas turbine engine according to the present invention showing a turbine section, Figure 2 is an enlarged cross-sectional view through the turbine section of the gas turbine engine shown in Figure 1 and Figure 3 is enlarged cross-sectional view through the turbine portion of the gas turbine engine of Figure 1 showing an alternative embodiment. A gas turbine engine turbomachine 10 is shown in Figure 1 and comprises in flow series an intake 12, a compressor section 14, a combustion section 16, a turbine section 18 and an exhaust nozzle 20. The compressor section 14 is arranged to be driven by the turbine section 18 and the compressor section 14 may comprise one or more compressor rotors (not shown) driven by one or more turbine rotors 26 via shafts (not shown). The gas turbine engine 10 operates quite conventionally in that air is compressed as it flows through the compressor section 14 to the combustion section 16. Fuel is injected into a combustion chamber 22 in the combustion section 16 and is burnt in the air supplied by the compressor section 14 to produce hot gases. The hot gases flow out of the combustion section 16 through the turbine section 18 and exhaust nozzle 20 to atmosphere. The hot exhaust gases drive the turbine section 18 which in turn drives the compressor section 14. The turbine section 18 includes a stage of high pressure nozzle guide vanes 24 between which the hot exhaust gases from combustion chamber 22 flow to the turbine section. A high pressure turbine motor 26 is located downstream of the high pressure nozzle guide vanes 24, and the high pressure turbine rotor 26 carries a plurality of circumferentially arranged, radially extending turbine blades 28. Downstream of the turbine blades 28 is located a stage of low pressure nozzle guide vanes 30, arranged circumferentially around and extending radially from a low pressure stator disc 32. A shroud 34 is arranged around and spaced from the tips of the high pressure turbine blades 28 by a clearance 36. The downstream end 38 of shroud 34 is cantilevered from the low pressure nozzle guide vanes 30 while its upstream end 40 of the shroud 34 is located between two radial stops, radial inner stop 42 and radial outer stop 44 on the high pressure nozzle guide vanes 24. A casing 46 encircles the nozzle guide vanes 24 and 34 and the shroud 34, and together with a member 48 defines an annular chamber 50 which receives cooling air via supply apertures 52. The member 48, located immediately surrounding the shroud 34, is provided with a plurality of holes 54 to allow air within the chamber 50 to be expelled in the direction of the shroud 34 to provide impingement cooling of the shroud 34. Used impingement cooling air flows through further apertures 56 in the shroud 34 into the gas stream flowing through the turbine section 18. The low pressure nozzle guide vanes 30 are provided with one or more passages 58 which extend radially through the low pressure nozzle guide vanes 30. These passages interconnect a chamber 60 defined between the casing 46 and the low pressure nozzle guide vanes 30 with a chamber 62 defined between the stator disc 32 and the low pressure nozzle guide vanes 30. The chamber 60 is connected to a bleed device 63 on the compressor section 14 of the gas turbine engine 10. The bleed device 63 comprises a plurality of apertures 64 in the casing 66 of the compressor section 14 thereby allowing air at a suitable pressure and temperature to be bled from the compressor section 14. This air is supplied to the chamber 60 via a pipe 68 and a valve 70. The passages 58 in the low pressure nozzle guide vanes 30 are preferably arranged to discharge the cooling air from the compressor section 14 such that it impinges directly upon the stator disc 32 for maximum cooling effect, ie at 90° to the stator disc surface. Alternatively the cooling air from the passages 58 may be directed to flow along the surface of the stator disc 32 in a radially inward direction. The valve 70 is a two-position on/off valve and preferably is operated by the engine or an aircraft control system. For example the rate of change of throttle, fuel flow, or other aircraft control demand is monitored and used to change the state of the valve 70. Alternatively the pilot could switch the position of the valve 70. In operation, for maximum fuel efficiency, the valve 70 is opened to allow cooling air to be bled from the compressor section 14 through the pipe 68 to the chamber 60. The cooling air then flows from chamber 60 through the passages 58 in the low pressure nozzle guide vanes 30 into the chamber 62. Cooling air discharged from the passages 58 is directed against, or over, the stator disc 32, cooling it and reducing its diameter, which in turn causes the low pressure nozzle guide vanes 30 to be moved radially inwardly. This radial inward movement of the low pressure nozzle guide vanes 30 causes the shroud 34 to move radially inwards thereby reducing the clearance 36 between the shroud 34 and the rotor blade 28 tips. The valve 70 is operated for example during cruise conditions of an aircraft when the speed of rotation of the engine remains relatively constant. For reducing the possibility of tip rubs between the turbine blade 28 tips and the shroud 34, the valve 70 is closed to prevent cooling air flowing from the compressor section 14 to cool the stator disc 32. The hotter stator disc 32 increases in diameter with corresponding radial outward movements of the low pressure nozzle guide vanes 30 and shroud 34 to increase the clearance 36 between the shroud 34 and the rotor blade 28 tips. The valve 70 is closed for example during transient conditions and violent manoeuvring of the aircraft. For example with a cooling flow of air to the arrangement shown in Figure 2 it is expected that this will provide a 0.4 mm reduction in diameter of the stator disc 32, a 0.5 mm reduction in clearance at the trailing edge of the turbine blade at point A, and a 0.59 mm reduction in clearance at the leading edge of the turbine blade at point B. This gives most movement at the leading edge of the turbine blade 28. In Figure 3 the shroud 34 is cantilevered from its upstream 40 on the high pressure nozzle guide vanes 24 giving most movement at the trailing edge of the turbine blade 28. The shroud may be cantilevered from the low pressure nozzle guide vanes or high pressure nozzle guide vanes by suspending the shroud from the mid section of a lever pivoted on the high pressure nozzle guide vanes or low pressure nozzle guide vanes. This would ensure that the clearance changed uniformly at all points between the shroud and rotor blades. Although the invention has been described with reference to gas turbine engines it may be equally applicable to other turbomachines. The invention may also be applicable to compressors of gas turbine engines, and it may be possible to supply hot air or fluid to the stator disc to increase the clearance to prevent rubs and not to supply hot air during cruise conditions.

Claims: A turbomachine comprising a rotor having a plurality of radially outward extending rotor blades (28), on one side of the rotor a stator assembly including a stator disc (32) and a plurality of stator vanes (30) arranged in a first stage extending radially outward from the stator disc (32), shroud means (34) encircling the rotor blades (28) at a radial clearance (A,B) therefrom, an opposite side of the shroud means (34) being located between the inner and outer radial stops on an adjacent stator assembly to limit radial movement of the shroud means (34), and a fluid supply means (60,68) controlled by valve means (70) and having at least one passage (58) extending radially through the stator assembly, characterised in that the shroud means (34) is cantilevered on one side from the first stage of stator vanes (30), and the at least one passage (58) is arranged to direct a supply of fluid to impinge directly onto the stator disc (32) and the valve means (70) is operated to control the supply of the fluid onto the disc (32) in order to influence the temperature and therefore the diameter of the disc (32) and thereby control the radial clearance (A,B) between the tips of the rotor blades (28) and the shroud means (34). A turbomachine as claimed in claim 1 in which the stator disc (32) and the first stage of stator vanes (30) are located downstream of the rotor blades (28). A turbomachine as claimed in any preceding claim in which the valve means (70) comprises an on/off valve arranged to supply cooling or heating fluid in a first mode of operation and to terminate the supply of cooling or heating fluid in a second mode of operation. A turbomachine as claimed in any preceding claim in which the supply of cooling or heating fluid is a supply of cool fluid to cool the stator disc (32), the valve means (70) is arranged to supply cool fluid to the stator disc (32) to cool the stator disc to reduce the clearance between the shroud (34) and the rotor blades (28) during cruise and the valve means (70) is arranged to terminate the supply of cooling fluid to the stator disc (32) to increase the clearance between the shroud and the rotor blades (28) during high speed operation of the turbomachine. A turbomachine as claimed in any preceding claim in which the turbomachine is a gas turbine engine. A turbomachine as claimed in claim 8 in which the rotor is a turbine rotor and the stator is a turbine stator. A turbomachine as claimed in claim 9 in which the means to supply cooling or heating fluid includes bleed means to bleed fluid from the compressor of the gas turbine engine.

Inventors: PELLOW TERENCE R, PELLOW, TERENCE R.
Applicants: ROLLS ROYCE PLC, ROLLS-ROYCE PLC
Application year: 1997
Main classification: F01D11
IPC codes: F01D11