CASCADED AIR SUPPLY FOR GAS TURBINE COOLING

Abstract: In a combustion turbine, the first stage turbine cooling air leaking through the torque tube seal 32 is diverted in a large part through the axial passage 52 by virtue of a metering opening 72 being provided in a downstream side plate, so that the diverted cool air is available to the second stage disc cavity for leakage and cooling.

Description: CASCADED AIR SUPPLY FOR GAS TURBINE COOLING This invention relates to a gas turbine, and in particular to gas turbine cooling. In gas turbines cooling air is supplied to the rotor disc areas to prevent rotor disc material overheating. For example, with a three-stage turbine and a 19 stage compressor, air is delivered to the third stage stator nozzle vanes from, say, the seventh stage of the compressor, the second stage nozzle vanes receive air from, say, the thirteenth stage of the compressor, and a third source of air, which is fully compressed and taken from the shell into which the compressor discharges, is cooled to a significantly lower temperature, and fed to a location on the torque tube bounded by dynamic seals from which it enters the rotor to cool the blades and inner portions of the discs or leaks through the dynamic seals. The turbine side dynamic seal leakage air serves in part to cool the first stage disc cavity, but its most important function is to prevent the flow of hot gases from the blade path upstream of the first stage blades from passing through a dynamic seal between the upstream side of the first stage blade assemblies and an opposing stationary part. In other words, in order to maintain a cool disc environment, it is desirable, if not necessary, to pass some air from the disc cavity areas into the blade path, rather than in an oppo ¯site direction. All three stiijarate sources of air referred to before perform this same function, as well as a cooling function in the disc cavity areas. Whereas the amount of air supplied to the second and third stages of the turbine is closely matched to perform the desired cooling and leakage functions, that which leaks through the turbine side torque tube seals exceeds the amount required to perform the first stage disc cooling and leakage function. Since excessive use of cooling flow reduces gas turbine efficiency it is desirable to reduce overall cooling air flow to a minimum. An object of this invention to provide a construction in a turbine of the type described to accomplish this.. According to the present invention, a gas turbine having a staged compressor and a staged turbine includes at least first and second stage rotor discs carrying indivdu- al rotor blade assemblies at the disc peripheries with an axially extending passages defined between adjacent blade assemblies, and a turbine cooling system in which cooling air is supplied through stator vane nozzles of at least said second stage of the turbine, and cooling air is also supplied through a torque tube seal to the disc cavity of said first stage disc, characterized by seal means between adjacent blade assemblies at the radially outer boundary of said axially extending passages, and plate means at the downstream end of said axial extending passage, said plate means including opening means therein sized, reLative to the effective opening area of the upstream dynamic seal between the blade assemblies and the opposing stationary part, to pass a substantional part of air from said first stage cavity to the next downstream stage cavity. Conventionally, the torque tube tube leakage air flow is divided after cooling the first disc cavity so that only a quantity approximating that to satisfy the requirement of leakage into the blade path is leaked therethrough, while the majority of the air is metered from an axial extending passage between the first stage blade assemblies into the second stage disc cavity where it performs a cooling function and additionally provides part or all of the air to satisfy the leakage function into the blade path. In the gas turbines provided by the assignee of this application this blade path leakage requirement is minimized by providing seal means between adjacent blade assemblies of the turbine first stage at the radially outer boundaries of the axially extending passages between the blades, and by providing plate means at the downstream ends of the axial extending passages. The "cascading" of the torque tube leakage air is obtained by providing opening means in this plate means sized, relative to the effective opening area of the upstream dynamic seal between the blade assembly and the opposing stationary parts and other minor leakage paths to pass the majority of air from the first stage cavity to the next downstream stage cavity. The invention will now be described, by way of example, with reference to the accompanying drawings in which: Figure 1 is a longitudinal sectional view, principally of the bladed portion, of an axial flow turbine having a cooling arrangement; Fig. 2 is a fragmentary side view of a rotor blade assembly with a seal plate partly broken and sectioned, in installed relation on the blade assembly; Fig. 3 is an isometric view of the fragmentary portion of the blade assembly, and with a fragmentary portion of the seal plate for the assembly shown in installed relation; and Fig. 4 is a sectional view corresponding to one taken along the line IV-IV, but also including a pin in the groove, and a portion of the abutting adjacent blade assembly. Fig. 1 shows a turbine portion of a gas turbine, a rotor 10 includes first, second, and third stage rotor discs 12, 14, 16, respectively, which carry on their circumferences the rotor blade assemblies 18, 20 and 22, respectivelyf eash:-of which has the bladed portion of the assembly in the hot gas path 24, which is sometimes called the blade path. Each stage of the turbine has its fixed stator vanes identified as 26, 28 and 30. In one type of turbine manufactured by the present applicants at least one dynamic seal 32 is provided between the stationary turbine structure and the rotating torque tube 34. In that type of turbine, air which has been fully compressed by the compressor (not shown) is taken from the combustor shell 36, passed through an external heat exchanger to cool the air and then is direct ed to the vicinity of the torque tube as. indicated by the solid line arrow at the upstream end of the seal 32. From there some air enters the torque tube through holes (not shown) and other air leaks through dynamic seal 32. While the air in the combustor shell may be at a temperature of 7500F (3990C) depending upon ambient temperatures, the air passing to the dynamic seal may be in the neighborhood of 4000F (2040C) after passing through the heat exchanger. Because of manufacturing tolerances and thermal expansion and contraction requirements, the labyrinth seal 32 is imperfect and air is expected to leak therethrough and into the disc cavity 38 of the first turbine stage as indicated by the arrow at the downstream end of the seal. This air performs a cooling function in the first disc cavity 38 and is then passed to the upstream side of the blade assembly 18 where, in the prior art turbine, all of it exits through the dynamic seal 40 at the upstream side of the blade assembly 18 or other minor leakage paths not shown. This leakage of the air through seal 40 prevents fluid in the hot gas or blade path 24 from passing into the rotor disc areas. Another source of cooling air for the turbine is air which is bled from some upstream stage, such as the thirteenth stage of a nineteen stage compressor in the particular embodiment being used as an example, and this air is then fed to the space 42 from which part of it leaks into the hot gas path to prevent reverse flow and part of it flows down into stator vane or nozzle 28, from which part of it performs a cooling function of the vane and leaks into the hot gas path and another part passes down into the second disc cavity space 44 to provide cooling and leakage for the second interstage seals 45 and into the gas stream at seal 47. The air which is bled from the thirteenth stage and passed to the space 42 may have a temperature in the order of 6000F (3160C) in accordance with ambient temperatures. In any event, with that particular turbine, the range of air temperatures for the second stage cooling is significantly greater than the temperature range of the air which passes through the seal 32 and to the first stage. Finally, additional air from another source, such as the seventh stage of the compressor is bled therefrom and passed to the space 46 for passage through the vane 30 to the third interstage seal cavity 48 for purposes of leakage. The temperature of this air may be in the range of 3500F (1770C), again depending upon ambient. Substantially all of this air is used for leakage from the third interstage disc cavity through seals 49 and 51, since the temperatures in the third stage of the turbine are significantly less than at the first stage, and substantially no cooling of the vane 30 is required to be carried out. In the prior art embodiment, a side plate 50 on the downstream side of each blade assembly is provided to seal as well as is reasonably possible against the leakage of the air from the first disc cavity 38 to the second disc cavity 44 so that, in the prior art embodiment, the air indicated as passing in the direction of the solid arrow 52 is absent. However, in the present invention, the majority of this air from the first disc cavity will flow as indicated by the directional arrow 52 to the second disc cavity, as will be explained in the description of the structure as shown in Figs. 2-4. As seen in Figs. 2 and 3, each first stage blade assembly 18 includes- the-lade 54 which extends into the blade path, the blade platform 56, the blade extension portion 58 and the root portion 60 which is of conventional serrated form and which is received in a complementary notched configuration at the periphery of the rotor disc. The blade assemblies are installed in the disc in the. conventional fashion with each end of the platform being in relatively close abutting relation to a facing end of a platform of an adjacent blade assembly. However, again because of required tolerances and expansion and contraction, there is some slight gap, such as 20 mils, for example, between the closely adjacent platforms. To substantially seal this gap, grooves 62 and 64 are provided in one end of the platform of each blade assembly, and pins (for example 66 in Fig. 4) are loosely disposed in the groove so that during operation, the pin will be moved by centrifugal force as indicated in Fig. 4 to substantially seal against the adjacent blade assembly 68. Since the blade extension portions 58 are recessed relative to the ends of the platform, and since the radially outer end of the rotor disc steeples (not shown) only extend radially outwardly sufficiently far to accommodate the serrations of the root 60, an axially extending passage is formed between the complementary recesses of the adjacent blade assemblies. It is through this generally axially extending space between adjacent blade assemblies, sealed at its radially outer boundaries by the centrifugally loaded pins, that cooling air can flow as indicated by the arrow 52. However, the blade assemblies do each have a downstream side plate, identified by the different numeral 70 in Figs. 2 and 3, since this side plate differs from that side plate 50 in Fig. 1 because of one important difference. That important difference is the provision of a metering hole 72 in an area of the side plate which is in communication with the axially extending passage. That hole is sized, relative to the effective opening area of the upstream dynamic seal formed at 40 and other minor 4- > eÅaRag-e pth not discussed sor that the--majo-rity of-t-he ain furnished to the first disc cavity will pass through the axially extending passage and out the opening 72 to the second stage disc cavity. As an example of numerical values, a typical turbine may pump 3 to 4 pounds of air per second through the labyrinth seal while about 1 pound of air per second through the dynamic seal 40 will guarantee an outflow into the blade path, as well as having provided some cooling of the first disc cavity. Thus in the preferred embodiment, the hole- 72 is sized to pass 2 to 3 pounds of air per second through the axial passage 52 and into the second stage disc cavity. As a result, the air which is bled from the compressor to feed the second stage of the turbine can be significantly reduced, such as in the order of 25% or more so that its principal function will be that of providing leakage air at the radially outer portion of the second stage, and cooling of the vane 28. Most or all of the requirement for Leakage air and cooling for the second disc cavity can be performed, in an arrangement according to the invention, by the air which passes through the axial passage 52. Thus, the torque tube leakage air is efficiently utilized to give an overall reduction in total cooling air consumption, thereby improving engine efficiency. While the invention has been described in connection with one particular type of turbine using different temperature cooling air sources because this is one of the best applications of the invention, the cascading principle can also be usefully applied to turbines having a common source for cooling air.

Claims: 8 CLAIMS: 1. A gas turbine having a staged compressor, and a staged turbine including at least first and. second stage rotor discs carrying individual rotor blade assemblies at the disc peripheries with an axially extending passages defined between adjacent blade assemblies, and a turbine cooling system in which cooling air is supplied through stator vane nozzles of at least said second stage of the turbine, and cooling air is also supplied through a torque tube seal to the disc cavity of said first stage disc, characterized by seal means between adjacent blade assemblies at the radially outer boundary of said axially extending passages, and plate means at the downstream end of said axial extending passage, said plate means including opening means therein sized, relative to the effective opening area of the upstream dynamic seal between the blade assemblies and the opposing stationary part, to pass a substantional part of air from said first stage cavity to the next downstream stage cavity. 2. A turbine as claimed in claim 1 wherein a substantial part of air passing to said next downstream stage cavity comprises in the order of 2 to 3 times the amount of air passed through said effective opening area of the upstream dynamic seal. 3. A turbine as claimed in claim 1 or 2 wherein said seal means comprises centrifugally loaded seal pins in groove means in each said blade assembly. 4. A turbine as claimed in any one of claims 1 to 3 wherein said cooling air supplied through said torque tube seal is of a lower temperature range than said cooling air supplied through said vane nozzles.

Inventors: HOLDEN, PAUL CLARENCE, SMED, JAN PEER
Applicants: WESTINGHOUSE ELECTRIC CORPORATION
Application year: 1984
Main classification: F01D5
IPC codes: F01D5