A gas turbine engine such as that used for powering an aircraft in flight includes a compressor for pressurizing ambient air which is mixed with fuel in a combustor and ignited for generating combustion gases which flow downstream through one or more turbines which extract energy therefrom. The hot combustion gases are channeled from the combustor through various flow confining structures including conventional turbine nozzles, turbine frames, and turbine rotor blade stages. All of these components are typically cooled from the adverse affects of the hot combustion gases by preferentially channeling cooling air bled from the compressor which returns to the combustion flowpath once it has performed its cooling function.
For example, a turbine frame is typically provided downstream of the combustor and includes a radially outer casing and a radially inner hub with a plurality of circumferentially spaced apart struts extending radially therebetween. The hub supports a conventional rotor shaft, with the reaction forces therefrom being channeled radially outwardly through the struts and into the casing. Since the struts extend through the combustion gas flowpath, they are typically surrounded by vanes or fairings which protect the struts from the combustion gases. The fairings are joined at their radially outer ends to an annular outer band or liner, and joined at their radially inner ends to an annular inner band or liner.
The outer band, for example, is typically spaced radially inwardly from the casing to define a generally open annular plenum which is suitably provided with bleed air from the compressor for cooling both the casing and the outer band in this vicinity. The cooling air channeled through the plenum absorbs heat at a generally uniform heat transfer rate along the various surfaces defining the plenum. However, the temperatures of the components defining the plenum vary considerably, with the hotter components enjoying less cooling than the cooler components which, therefore, creates undesirable thermal gradients through the various components with attendant thermally induced stresses.
Furthermore, the hotter components will radiate heat to adjacent components which requires even more cooling thereof. In order to reduce the heating effect from the hotter components, conventional sheet metal heat, or thermal radiation, shields are typically provided where necessary between a hot component and a cold component to interrupt the radiation of heat therebetween for reducing component temperatures and thermal gradients therein.
However, although a radiation shield can effectively block thermal radiation heating of cold components in the engine, thermal gradients typically remain especially in the axial direction which is the direction of combustion gas flow. For example, in the turbine frame described above, the fairings and inner and outer bands begin upstream of the struts and extend downstream therefrom for a considerable axial length. And, cooling air is typically channeled into the plenum between the outer band and the casing near its center and then fills the plenum. Since cooling is more effective where the cooling air is introduced into the plenum than at the axially forward and aft ends of the plenum, the turbine casing runs hotter at its forward and aft ends as compared to its mid-portion which creates axial thermal temperature gradients and corresponding thermal stresses therein. And, cooling air picks up heat between where it is introduced into the plenum and its eventual exit into the flowpath. This creates circumferential thermal gradients, out-of-round distortion of the case, increased turbine clearances, and performance loss.
USPTO Backgrounds
USPTO Backgrounds