Abstract:
In a turbine for aeronautic engines, a first and at least a second turbine disk rotating around a common axis respectively carrying a first and a second moving-blade crown, with which first and, respectively, second axial passages are defined through which a cooling air mass for the turbine disks can pass, a device for conveying cooling air being interposed between the first and the second axial passages to receive the mass of air passing through the first axial passages and send it through the second axial passages.

Description:
The present invention relates to a gas turbine for aeronautic engines. 
     BACKGROUND OF THE INVENTION 
     As is known, a gas turbine for aeronautic engines generally comprises a number of rotating bladed sectors, each of which, in turn, comprises a turbine disk connected to adjacent turbine disks and carrying a coupled blade crown. 
     As is also known, turbine disks are components that are subjected to high stress, both mechanical, due to the effect of centrifugal components and, above all, thermal, since they operate in an extremely high temperature environment due to close vicinity with the flow of hot gases that impact the blades. For optimal turbine operation it therefore becomes necessary to control the operating temperature of these disks, maintaining the operating temperature below a set or critical threshold value. 
     To that end, it is known to send to each of the turbine disks its own cooling airflow, separate from the other cooling airflows. Each cooling airflow is normally formed by bleeding a predetermined quantity of air from the compressor and conveying the bled air to the area of connection of the blades to the respective turbine disk. In the area of connection of the blades to the disk, the air is made to flow through passages, each one being defined on one side by a slot in the turbine disk to be cooled and by the leading portion or lobe of the relevant blade, on the other. While traversing the passages, the cooling air progressively heats up, carrying away heat by convection; at the exit, the heated air is first fed into a mixing chamber where it mixes with part of the mentioned flow of hot gases, forming a mixture of lower temperature that passes over the side walls of the blade and the turbine disk, after which the same mixture is reinserted in the flow of hot gases before this flow passes over the bladed sector arranged downstream of the cooled turbine disk. 
     Although it is used, for various reasons, the described cooling method is found to be less than satisfactory. 
     First of all, cooling of the disks is performed in conditions of low efficiency and therefore the cooling capacity of the incoming air is only exploited in part. In consequence, the air exiting from the respective turbine disks has a relatively low temperature for which, when mixed with the hot gases entering the downstream stage, it significantly lowers the temperature of the hot gases in an undesired manner. 
     With bleeding being carried out for each turbine disk, as the number of turbine disks increases, so does the quantity of air that is used and the overall efficiency consequently decreases. 
     SUMMARY OF THE INVENTION 
     The object of the present invention is that of making a gas turbine for aeronautic engines, the embodying characteristics of which enable the above described problems to be resolved in a simple and inexpensive manner. 
     According to the present invention a gas turbine for aeronautic engines is produced comprising a first and at least a second rotating bladed sector respectively comprising a first and a second turbine disk arranged coaxially to an axis of the turbine and respectively carrying a coupled first and second moving-blade crown, the first and the second turbine disk defining with the respective blades first and, respectively, second passages through which a cooling air mass for said turbine disks can pass, characterized in that it further comprises means for conveying said cooling air interposed between said first and second bladed sector to convey the cooling air exiting said first passages towards said second passages. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWING 
       The invention will now be described with reference to the attached FIGURE, which partially illustrate a preferred non-limitative embodiment of a gas turbine for aeronautic engines made according to the principles of the present invention. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     In the attached FIGURE, reference numeral  1  indicates, as a whole, a gas turbine for an aeronautic engine (not shown). The turbine  1  comprises a plurality of rotating bladed sectors  2 , only two of which are visible in the attached FIGURE and a plurality of stator bodies  3  only two of which are visible in the attached FIGURE, arranged between the two bladed sectors  2 . 
     The rotating bladed sectors  2  extend coaxially to a turbine axis, indicated by reference numeral  5 , and respectively comprise turbine disks  6  and  7 , rotating around axis  5  and, in turn, each comprising a respective disk-like central portion  8  and  9  and, for each disk-like central portion, an associated pair of lateral, internally-flanged tubular bodies  10  and  11 . 
     The lateral tubular bodies  10  and  11  integrally protrude from the associated disk-like central portion  8  and  9  in opposite directions and terminate with respective inner flanges  10   a  and  11   a  stably connected to one another by a ring of screws  13 , only one of which is visible in the attached FIGURE. 
     The turbine disks  6  and  7  carry respective blade crowns, respectively indicated by reference numerals  15  and  16  and coupled in a known manner, which in turn comprise respective roots  15   a  and  16   a  stably coupled to the associated disk-like central portion  8  and  9  and associated outer shaped portions  15   b  and  16   b  over which a functional flow A of hot gases passes during use. 
     Each root  15   a  is housed in an associated seat made on the outer perimeter of the disk-like central portion  8  and, with the associated seat, defines a respective through opening or axial passage  18 , parallel to axis  5  in the described example. According to a variant that is not shown, the axial passage  18  is inclined with respect to axis  5 . 
     Each root  16   a  is instead housed in an associated seat made on the inner perimeter of the disk-like central portion  9  and, with the associated seat, defines a further through opening or axial passage  19 , always parallel to axis  5 . 
     Always with reference to the attached FIGURE, the passages  18  and  19  constitute part of a closed circulation circuit  20  of a cooling airflow for the turbine disks  6  and  7 . In addition to the passages  18  and  19 , the circuit  20  comprises a conveying device  22  to receive the cooling air mass exiting passages  18  and convey this mass of air to the inlet of and through passages  19 . 
     In the particular example described, the device  22  comprises two shaped annular bodies, indicated with reference numerals  23  and  24 , which respectively surround portion  10  and portion  11  and have respective inner flanges  23   a  and  24   a  arranged in contact with each other and tightened in a pack between flanges  10   a  and  11   a  by screws  13 . The annular bodies  23  and  24  comprise respective intermediate portions  23   b  and  24   b  with a rectilinear generatrix, which extend from the associated inner flanges  23   a  and  24   a  to the respective blade crowns and, with portions  10  and  11 , define two sections  25  and  26  of a duct  27 . Sections  25  and  26  communicate with each other through a crown of openings D made through flanges  23   a  and  24   a.    
     Finally, always with reference to the attached FIGURE, the annular bodies  23  and  24  comprise respective terminal portions  23   c  and  24   c  that, in the particular example described, have inner diameters greater than those of the intermediate portions  23   b  and  24   b , are coupled in a substantially fluid-tight manner to the associated blade crowns  16  and  15  and, together with blade crowns  16  and  15  and the associated turbine disk  7  and  8 , define respective chambers  29  and  30 . Chambers  29  and  30  constitute part of circuit  20  and communicate with the associated sections  25  and  26  of duct  27  on one side and with the respective openings  19  and  18  on the other. In this way, cooling air passing through the openings  18  is collected in chamber  30  and from here sent through duct  27  to chamber  29 ; this air proceeds from chamber  29 , passing through passages  19 , in this way also cooling turbine disk  7 . 
     Still with reference to the attached FIGURE, each portion  15   b  is coupled to an adjacent stator body  3  by means of a respective controlled-leakage seal  32 . In the particular example described, the seal  32  comprises a pair of annular fins  33  and  34 , which are carried by portion  15   b  and in which fin  33  cooperates with a body  35  of abradable material carried by the stator body  3 , while fin  34  cooperates directly with an inner surface of the stator body  3 . In normal running conditions, fins  33  and  34  define a passage  36  through which a precise part C of the flow A of hot gases is drawn. 
     In the described example, each portion  16   b  is instead coupled to the adjacent stator body  3  by means of an associated seal  37 , which is similar to seal  32  and defines a reinjection passage  38  for the bled part C of the hot gas back into the flow A of hot gases. 
     The cooling air for the turbine disks  6  and  7  transiting inside circuit  20  is insulated from flow A of hot gases, but above all from part C bled through seal  32 , by a heat barrier, indicated as a whole by reference numeral  40 . 
     The heat barrier  40  comprises a mechanical guide barrier  41  stably connected to the stator bodies  3  and defining with these stator bodies  3  an additional annular feed duct  42 , which is able to receive part C of the hot gas exiting seal  32  and convey this part C of hot gas to seal  38 , which permits its reinjection into flow A of hot gases. 
     Always regarding the particular example described, duct  42  houses the seals  32  and  38  inside its axially ending terminal sections, of different volumes, while the mechanical barrier  41  comprises two annular metal bodies  43  and  44 , which are coupled to each other in a fluid-tight manner and protrude in axially opposite directions from a perforated support appendage  45  obtained near the junction area of the turbine disks  6  and  7 . 
     In addition to the mechanical barrier  41 , the heat barrier  40  also comprises an annular chamber  47  arranged between duct  42  for conveying part C of the bled hot gas and duct  27  for conveying the cooling air. Chamber  47  is circumferentially delimited by bodies  43  and  44  on the outer side and by bodies  23  and  24  on the inner side and communicates with an inlet and with an outlet of duct  42  through two throttled passages, indicated by reference numerals  47   a  and  47   b . In use, chamber  47  houses a mass of air that, due to the shape of passages  47   a  and  47   b , in practice can only move in the circumferential direction and thus defines a thermally insulating cushion that separates the hot flows A and C from the flow B of cooling air, preventing the latter from being heated up during transit from one turbine disk to the next. 
     First of all, from the foregoing it is evident that in the described turbine  1  a single cooling airflow common to all the turbine disks  6  and  7  is used. In fact, each circuit  20  enables taking the cooling air exiting an upstream turbine disk and sending it to a turbine disk arranged immediately downstream. In this way, always with respect to known solutions, the mass of air destined to cooling the turbine disks is significantly reduced. 
     The same cooling air is never mixed or added to the flow of hot gases and therefore the temperature of these hot gases is not affected by the mass and temperature of the cooling air. 
     Furthermore, in order to avoid undesired heating of the air during transfer from one turbine disk to the next, a thermally insulating barrier is provided in the described turbine  1  that, in the particular example described, comprises a mechanical barrier, with the function of reinjecting the bled hot gases into the main flow again, and an insulating air cushion to thermally separate the mechanical barrier of the device provided to guide the cooling airflow through the various turbine disks. 
     From the foregoing, it is apparent that changes and modifications may be made to the turbine  1  described herein without leaving the scope of protection defined in the independent claims. 
     In particular, the device to guide the cooling air towards the axial passages of the downstream turbine disk could be constructively different from that described by way of example. In particular, the cooling air could advance along a defined path that is not closed due, for example, to possible bleeding in the interface area between portions  23   c  and  24   c  and blades  15  and  16 .