Active turbine tip clearance control system

#The described subject matter relates generally to a gas turbine engine and more particularly, to an active tip clearance control (ATCC) system of a gas turbine engine.

#Conventional active tip clearance control (ATCC) systems in a turbofan gas turbine engine use the fan driven bypass air, a portion of which is directed through a diffusion duct to the high pressure turbine case manifold. This portion of bypass air is directed to flow over the high pressure turbine case through a series of impingement holes. A modulating valve may be incorporated between an air inlet and the diffusion duct. This valve adjusts the portion of bypass air flow according to the engine requirements such that the appropriate tip clearance between the turbine blades and the turbine shroud is maintained. Conventional ATCC systems rely on the pressure differential between an inlet scoop at the bypass duct and a downstream location of the bypass duct where the vent cooling air is dumped. For engines in which this pressure differential is very low, the operation and/or efficiency of the ATCC system may be at risk.

#Therefore, there is a need for improvements to an ATCC system.

#Reference is now made to the accompanying drawings depicting aspects of the described subject matter, in which:

#FIG. 1 is a schematic cross-sectional view of a turbofan gas turbine engine having an active tip clearance control (ATCC) system;

#FIG. 2 is a schematic cross-sectional view of the turbofan gas turbine engine of FIG. 1, showing the ATCC system according to one embodiment;

#FIG. 1 illustrates a turbofan gas turbine aircraft engine presented as an example of the application of the described concept, including a housing or nacelle annular outer case 10, an annular core case 13, a low pressure spool assembly seen generally at 12 which includes a fan assembly 14, a low pressure compressor assembly 16 and a low pressure turbine assembly 18, and a high pressure spool assembly seen generally at 20 which includes a high pressure compressor assembly 22 and a high pressure turbine assembly 24. The annular core case 13 surrounds the low and high pressure spool assemblies 12 and 20 in order to define a main fluid path (not numbered) therethrough. In the main fluid path there is provided a combustor to constitute a gas generator section 26. An annular bypass air duct 28 is defined radially between the annular outer case 10 and the annular core case 13 for directing a main bypass air flow (not numbered) driven by the fan assembly 14, to pass therethrough and to be discharged to the atmosphere to create a bypass air thrust to the aircraft engine.

#Referring to FIGS. 1-2 and 5, the turbofan gas turbine aircraft engine according to one embodiment includes an active tip clearance control (ATCC) system 30 located within the core case 13. The ATCC system 30 has an inlet passage 32 connected to the bypass air duct 28 at a location for example, immediately downstream of the fan assembly 14. An inlet 34 of the inlet passage 32 may be defined in the core case 13 or may be defined as a scoop (not shown) incorporated with a radial hollow strut (not numbered) in the annular bypass air duct 28, for introducing a portion (indicated by arrows 31) of the bypass air flow into the ATCC system 30. A vent passage 36 is provided to the ATCC system 30 and is in fluid communication with the atmosphere, for example via the bypass air duct 28, as shown in FIG. 2. The vent passage 36 according to this embodiment, has an outlet 38 defined in the bypass duct 28 at a location downstream of the location of the inlet 34. The ATCC system 30 is sealed to prevent leakage of the portion 31 of the bypass air flow passing through the ATCC system 30. Therefore, the portion 31 of the bypass air flow passes through the ATCC system 30 is discharged only through the vent passage 36 to the bypass duct 28 and then to outside of the engine.

#The ATCC system 30 may include an annular manifold 40, which is positioned around the turbine assembly (either the high pressure turbine assembly 24 or low pressure turbine assembly 18), for example, around an annular turbine case 42 such as a turbine support case or a turbine shroud. The manifold 40 defines a annular plenum 44 therein and is connected with the inlet passage 32. Therefore, the portion 31 of the bypass air flow is introduced from the annular bypass air duct 28 through the inlet 34 and inlet passage 32 and then into the annular plenum 44. The manifold 40 may further include a shield 46 which is configured to contour an outer surface of the turbine case 42 and includes a plurality of holes 48 defined in the shield 46 (see FIG. 5), to allow the portion 31 of the bypass air flow to be discharged from the holes 48 and to impinge on the outer surface of the annular turbine case 42 in order to cool the annular turbine case 42 and other turbine components (not shown) which are directly connected to the turbine case 42, thereby reducing blade tip clearances.