Patent Abstract:
A gas turbine engine assembly includes a housing including an annular duct wall that at least partially defines a mainstream hot gas flow path; a stator assembly with a stator vane extending into the mainstream gas flow; and a turbine rotor assembly upstream of the stator assembly and defining a turbine cavity with the stator assembly. The turbine rotor assembly includes a rotor disk having a forward side and an aft side, a rotor platform positioned on a periphery of the rotor disk, the rotor platform defining an aft flow discourager, a rotor blade mounted on the rotor platform extending into the mainstream gas flow, and an aft seal plate mounted on the aft side of the rotor disk. The aft seal plate has a radius such that the aft seal plate protects the rotor disk from hot gas ingestion.

Full Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     This invention was made with Government support under Contract No. W911W6-08-2-0001 awarded by the US Army. The Government has certain rights in this invention. 
    
    
     TECHNICAL FIELD 
     The present invention generally relates to gas turbine engine assemblies, and more particularly relates to turbine assemblies with improved cooling characteristics. 
     BACKGROUND 
     A gas turbine engine may be used to power various types of vehicles and systems. A particular type of gas turbine engine that may be used to power aircraft is a turbofan gas turbine engine. A turbofan gas turbine engine conventionally includes, for example, five major sections: a fan section, a compressor section, a combustor section, a turbine section, and an exhaust section. The fan section is typically positioned at the inlet section of the engine and includes a fan that induces air from the surrounding environment into the engine and accelerates a fraction of this air toward the compressor section. The remaining fraction of air induced into the fan section is accelerated into and through a bypass plenum and out the exhaust section. 
     The compressor section raises the pressure of the air it receives from the fan section, and the resulting compressed air then enters the combustor section, where a ring of fuel nozzles injects a steady stream of fuel into a combustion chamber formed between inner and outer liners. The fuel and air mixture is ignited to form combustion gases, which drive rotors in the turbine section for power extraction. The gases then exit the engine at the exhaust section. 
     In a typical configuration, the turbine section includes rows of stator vanes and rotor blades disposed in an alternating sequence along the axial length of a generally annular hot gas flow path. The rotor blades are mounted at the periphery of one or more rotor disks that are coupled in turn to a main engine shaft. 
     In most gas turbine engine applications, it is desirable to regulate the operating temperature of certain engine components in order to prevent overheating and potential mechanical failures attributable thereto. As such, most turbine components, particularly the stator vane and rotor blade assemblies may benefit from temperature management in view of the high temperature environment of the mainstream hot gas flow path. Accordingly, in many turbine sections, the volumetric space disposed radially inwardly or internally from the hot gas flow path includes an internal cavity through which a cooling air flow is provided. The cooling of the internal engine cavity attempts to maintain the temperatures of the rotor disks and other internal engine components that are suitable for their material and stress level. 
     However, in many conventional engines, relatively high levels of cooling air flows have been used to obtain satisfactory temperature control of the components within the internal engine cavity. In addition, the demand for cooling flow may be impacted by an irregular and unpredictable ingestion of mainstream hot gases from the hot gas flow path into the internal engine cavity. Various attempts to prevent hot gas ingestion between adjacent stator vanes and rotor blades have primarily involved the use of overlapping lip-type structures in close running clearance, often referred to as flow discouragers, but these structures have not been as effective as desired. Moreover, it is generally desirable to employ mechanisms to minimize this cooling air since air from the compressor used for cooling is not available for combustion. Additionally, temperature control of the flow discouragers should also be considered. If the flow discouragers are exposed to undesirably high temperatures, they may deform, which may impact their primary functions. 
     Accordingly, it is desirable to provide an improved gas turbine engine assembly that maintains proper temperature control. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention. 
     BRIEF SUMMARY 
     In accordance with an exemplary embodiment, a gas turbine engine assembly includes a housing including an annular duct wall that at least partially defines a mainstream hot gas flow path configured to receive mainstream hot gas flow; a stator assembly comprising a stator vane extending into the mainstream gas flow; and a turbine rotor assembly upstream of the stator assembly and defining a turbine cavity with the stator assembly. The turbine rotor assembly includes a rotor disk having a forward side and an aft side, a rotor platform positioned on a periphery of the rotor disk, the rotor platform defining an aft flow discourager, a rotor blade mounted on the rotor platform extending into the mainstream gas flow, and an aft seal plate mounted on the aft side of the rotor disk. The aft seal plate has a radius such that the aft seal plate protects the rotor platform from hot gas ingestion of the mainstream hot gas flow path into the turbine cavity. 
     In accordance with another exemplary embodiment, a turbine assembly is provided for a gas turbine engine assembly defining a mainstream hot gas flow path that receives mainstream hot gas flow. The assembly includes a rotor disk having a forward side, an aft side, and a circumferential periphery, a rotor platform positioned on the periphery of the rotor disk, the rotor platform defining an aft flow discourager, a rotor blade mounted on the rotor platform extending into the mainstream gas flow, and an aft seal plate mounted on the aft side of the rotor disk. The aft seal plate defines at least one cooling channel configured to deliver cooling flow to the aft flow discourager. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The present invention will hereinafter be described in conjunction with the following drawing figures, wherein like numerals denote like elements, and wherein: 
         FIG. 1  is a partial cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment; 
         FIG. 2  is a partial cross-sectional view of a turbine section of the gas turbine engine of  FIG. 1  in accordance with an exemplary embodiment; and 
         FIG. 3  is an enlarged cross-sectional view of a portion of the turbine section of  FIG. 2 . 
     
    
    
     DETAILED DESCRIPTION 
     The following detailed description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. Furthermore, there is no intention to be bound by any theory presented in the preceding background or the following detailed description. 
     Broadly, exemplary embodiments discussed herein include gas turbine engine assemblies that maintain suitable temperatures and reduce or eliminate of the effects of hot gas ingestion. More particularly, exemplary gas turbine engine assemblies include a turbine rotor assembly with an aft flow discourager. An aft seal plate may be configured to cooperate with the aft flow discourager to protect the rotor disk components, including the aft flow discourager, from elevated temperatures and conditions. Additionally, the aft seal plate may have channels that deliver impingement cooling flow to the aft flow discourager. 
       FIG. 1  is a cross-sectional view of a gas turbine engine  100 , according to an exemplary embodiment. In general, exemplary embodiments discussed herein may be applicable to any type of engines, including turboshaft engines. The gas turbine engine  100  can form part of, for example, an auxiliary power unit for an aircraft or a propulsion system for an aircraft. The gas turbine engine  100  has an overall construction and operation that is generally understood by persons skilled in the art. The gas turbine engine  100  may be disposed in an engine case  110  and may include a fan section  120 , a compressor section  130 , a combustion section  140 , a turbine section  150 , and an exhaust section  160 . The fan section  120  may include a fan, which draws in and accelerates air. A fraction of the accelerated air exhausted from the fan section  120  is directed through a bypass section  170  to provide a forward thrust. The remaining fraction of air exhausted from the fan is directed into the compressor section  130 . 
     The compressor section  130  may include a series of compressors that raise the pressure of the air directed into it from the fan. The compressors may direct the compressed air into the combustion section  140 . In the combustion section  140 , the high pressure air is mixed with fuel and combusted. The combusted air is then directed into the turbine section  150 . 
     As described in further detail below, the turbine section  150  may include a series of rotor and stator assemblies disposed in axial flow series. The combusted air from the combustion section  140  expands through the rotor and stator assemblies and causes the rotor assemblies to rotate a main engine shaft for energy extraction. The air is then exhausted through a propulsion nozzle disposed in the exhaust section  160  to provide additional forward thrust. 
       FIG. 2  is a partial cross-sectional view of a turbine assembly such as the turbine section  150  of the gas turbine engine  100  of  FIG. 1  in accordance with an exemplary embodiment. In general terms, the turbine section  150  includes a mainstream flow path  210  defined in part by an annular duct wall  212  for receiving mainstream hot gas flow  214  from the combustion section  140  ( FIG. 1 ). 
     The turbine section  150  includes an alternating sequence of stator assemblies  220 ,  230  and rotor assemblies  240 . In the view of  FIG. 3 , first and second stator assemblies  220 ,  230  and one rotor assembly  240  are shown. The first and second stator assemblies will be referred to as “forward” and “aft” stator assemblies based on their relative orientation with respect to the illustrated rotor assembly  240 . In general, any number of stator and rotor assemblies  220 ,  230 ,  240  may be provided. As discussed in greater detail below, the mainstream hot gas flow  214  flows past the stator and rotor assemblies  220 ,  230 ,  240 . 
     The forward stator assembly  220  is formed by stator vanes  224  extending radially outward from a platform  226  to the wall  212 , and the aft stator assembly  230  is similarly formed by stator vanes  234  extending radially outward from a platform  236  to the wall  212 . The platforms  226 ,  236  can be directly mounted to the combustor (not shown), or coupled to the combustor through intervening components, to form a portion of the mainstream flow path  210  with the wall  212 . 
     The rotor assembly  240  is formed by turbine rotor blades  242  projecting radially outwardly from a circumferential rotor platform  244  mounted on the periphery of a rotor disk  246 , which in turn circumscribes a main engine shaft (not shown). During operation, the mainstream hot gas flow  214  drives the rotor blades  242  and the associated rotor assembly  240  for power extraction, while the stator assemblies  220  are generally stationary. 
     Turbine rotor cavities  250 ,  270  are formed between the stator assemblies  220 ,  230  and the rotor assembly  240 . In the depicted embodiment, the disk cavities  250 ,  270  will be referred to as a forward rotor cavity  250  and an aft rotor cavity  270  based on the position of the rotor assembly  240 . A forward gap  252  is formed between the mainstream flow path  210  and the forward rotor cavity  250 , and an aft gap  272  is formed between the mainstream flow path  210  and the aft rotor cavity  270 . As discussed in further detail below, a portion of the mainstream hot gas flow  214  may attempt to flow through the gaps  252 ,  272  during operation. If unaddressed, the elevated temperatures of the mainstream hot gas flow  214  may adversely affect certain components in the rotor cavities  250 ,  270 . 
     Various mechanisms of the turbine section  150  attempt to prevent, reduce, or mitigate the effects of the mainstream gas ingestion. For example, in the depicted exemplary embodiment, the forward gap  252  is defined by a stationary flow discourager  228  extending downstream from the platform  226  of the stator assembly  220  and a forward rotor flow discourager  254  extending upstream from the turbine platform  244 . Generally, the stationary flow discourager  228  and the forward rotor flow discourager  254  overlap one another such that the mainstream hot gas flow  214  flows over the discouragers  228 ,  254  and stays in the mainstream flow path  210  instead of flowing through the forward gap  252  into the forward rotor cavity  250 . Similarly, the forward gap  272  is defined by an aft rotor flow discourager  248  extending downstream from the platform  244  of the rotor assembly  240  and a forward stationary flow discourager  238  extending upstream from the stator platform  236 . Generally, the aft rotor flow discourager  248  and the stationary flow discourager  238  overlap one another such that the mainstream hot gas flow  214  flows over the flow discouragers  248 ,  238  and stays in the mainstream flow path  210  instead of flowing through the aft gap  272  into the aft rotor cavity  270 . 
     The rotor assembly  240  further includes a forward seal plate  256  that is generally concentric with the rotor disk  246  and is mounted on and rotates with a forward face of the rotor disk  246 . The forward seal plate  256  generally has a radius such that a peripheral portion  258  extends adjacent to the forward rotor flow discourager  254 . The forward seal plate  256  may form a forward seal plate cavity  260  with the forward face of the rotor disk  246 . The forward seal plate  256  cooperates with the stationary flow discourager  228  and forward rotor flow discourager  254  to prevent or inhibit hot gas ingestion. As discussed in greater detail below, the forward seal plate  256  also directs cooling air into the rotor disk  246 . 
     The rotor assembly  240  further includes an aft seal plate  276  that is generally concentric with the rotor disk  246  and is mounted on and rotates with an aft face of the rotor disk  246 . The aft seal plate  276  generally has a radius such that a peripheral portion  278  extends adjacent to the aft rotor flow discourager  248 . The aft seal plate  276  may form an aft seal plate cavity  280  with the aft face of the rotor disk  246 . As discussed in greater detail below, the aft seal plate  276  cooperates with the aft rotor flow discourager  248  and stationary flow discourager  238  to prevent, inhibit, or mitigate the effects of hot gas ingestion. 
     Additional temperature control mechanisms include cooling air  290  that flows through the rotor cavities  250 ,  270  and through the rotor assembly  240 . In particular, the cooling air  290  may be obtained as bleed flow from a compressor or compressor section  130  ( FIG. 1 ) and flows to the forward seal plate cavity  260  to assist in maintaining an appropriate temperature of the rotor disk  246  and forward seal plate  256 . The cooling air  290  may additionally flow through a disk channel  262  in the rotor disk  246 . A seal  296  may be provided between the forward seal plate  256  and the rotor disk  246  to minimize leakage between the aft seal plate cavity  280  and the disk channel  262 . The disk channel  262  may be in fluid communication with internal passageways (not shown) through the rotor platform  244  and within the rotor blade  242 . As such, during operation, the cooling air  290  is drawn through the rotor disk  246  and rotor blade  242  for cooling these components. In one embodiment, the cooling air  290  may form a cooling film on the surface of the rotor blade  242 . 
     The cooling air  290  may additionally flow from the disk channel  262  to the aft seal plate cavity  280  to assist in maintaining an appropriate temperature of the rotor disk  246  and aft seal plate  276 . As discussed in further detail below with reference to  FIG. 3 , the aft seal plate  276  defines a number of impingement cooling channels  292  that extend in a radial direction from the aft seal plate cavity  280 . In general, a number of impingement cooling channels  292  may be arranged circumferentially around the aft seal plate  276 . The impingement cooling channels  292  deliver the cooling air  290  to the underside of the platform  244 , particularly the aft rotor flow discourager  248 . A seal  294  may be provided to prevent leakage of the cooling air  290  and encourage flow into the impingement cooling channels  292 . In further embodiments, the impingement cooling channels  292  may receive cooling air  290  directly from the disk channel  262  or an alternate source. 
       FIG. 3  is an enlarged cross-sectional view of a portion  300  of the turbine section  150  of  FIG. 2 . In particular,  FIG. 3  illustrates the aft rotor flow discourager  248 , the peripheral portion of the aft seal plate  276 , and the impingement cooling channels  292 . 
     As noted above, ingested gas from the mainstream hot gas flow  214  may attempt to flow through the aft gap  272  into the aft rotor cavity  270  or through the forward gap  252  and the rotor assembly  240  to the underside of the aft flow discourager  248 . The aft seal plate  276  generally has an extended radius such that the peripheral portion  278  extends adjacent to the aft rotor flow discourager  248 . The aft seal plate  276  generally prevents, inhibits, or mitigates the effects of hot gas ingestion in this area by limiting the exposure of the rotor disk  246 , such as a majority or substantially all of the rotor disk  246 . The extended aft seal plate  276  may also limit hot gas flowing through the forward gap  252  to the underside of the aft flow discourager  248 . In general, the aft seal plate  276  is tucked under the aft flow discourager  248  as close as possible with consideration for manufacturing tolerances and relative radial deflections. For example, the aft seal plate  276  may have a radius that is at least 50% of the radius of the rotor disk  242 . In other exemplary embodiments, the aft seal plate  276  may have a radius that is at least 90%, 95%, or 100% of the radius of the rotor disk  242 . In one exemplary embodiment, the impingement gap (i.e., the gap between the aft seal plate  276  and the aft flow discourager  248 ) may be any suitable distance corresponding to the radius ratios discussed above. In other embodiments, the impingement gap may be a function of the diameter of the cooling channels  292 . For example, the ratio of the impingement gap and the diameter of the cooling channel  292  may be about 2:1. In other embodiments, the ratio may be any suitable ratio, including about 1:1 to about 1:3. In conventional turbine assemblies, the aft seal plate does not extend to adjacent the turbine flow discourager. 
     As also noted above, the impingement cooling channels  292  deliver cooling air  290  that directly impinges upon and cools the aft rotor flow discourager  248 . In conventional turbine assemblies, temperature control of the aft flow discourager is typically unaddressed, and as such, the aft flow discourager tends deform, particularly in a radially outward direction, which widens the gap and adversely affects the function of the flow discouragers. In general, the impingement cooling channels  292  are oriented such that the cooling air  290  strikes the aft rotor flow discourager  248  at an angle of approximately 90°, although other angles may be possible based on structural design and cooling requirements. In generally, the aft rotor flow discourager  248  is maintained at a temperature and stress combination such that little or no deformation of the discourager may occur. In general, the impingement cooling channels  292  may have a length/diameter ratio of approximately 2:1, although other ratios are possible such that satisfactory jets of cooling air  290  are established. In the depicted exemplary embodiment, the impingement cooling channel  292  extends past the seal  296  and cooling air is supplied from radially inward (i.e., below) the seal  296 . In further exemplary embodiments, the cooling channel  292  does not extend past the seal  296  and cooling air is supplied via controlled leakage past the seal  292 . The cooling air  290  from the impingement cooling channels  292  may also function as an ingestion inhibiting dynamic jet that assists in recirculating any ingested gas back into the mainstream flow path  210 . In some embodiments, the impingement cooling channels  292  may enable the aft flow discourager  248  to be extended and/or the stationary flow discourager  238  to be shortened relative to conventional assemblies. In other embodiments, the lengths of the aft flow discourager  248  and the stationary flow discourager  238  are not modified. 
     Computational fluid dynamics (CFD) analysis may be used to determine the number, orientation, dimension, and position of the impingement cooling channels  292 . In general, design of impingement cooling channels  292  may depend on factors including application and engine design. In one exemplary embodiment, the impingement cooling channels  292  are provided to maintain the aft flow discourager  248  to a suitable temperature. Considerations may include engine application, required heat extraction, stress analysis, the temperature and pressure of the cooling air, and convective cooling effectiveness. The impingement cooling channels  292  may be formed, for example, by EDM or STEM drilling. 
     The aft seal plate  276  may additionally include one or more axial flanges  282 ,  284  that provide additional support to the aft seal plate  276  during operation. Particularly, axial flanges  282 ,  284  are configured such that undesirable deflections do not occur as the aft seal plate  276  rotates. For example, axial flange  282  may prevent the seal  294  from separating from the aft face of the rotor disk. Similarly, axial flange  284  maintains the position of the impingement cooling channels  292  relative to the aft rotor flow discourager  248 , which may be important if the peripheral portion  278  has a reduced amount of material resulting from the formation of the impingement cooling channels  292 . 
     Accordingly, exemplary embodiments provide a turbine section  150  with improved temperature control characteristics. In general, in combination or individually, the extended radius aft seal plate  276  and the impingement cooling channels  292  may mitigate and/or protect the aft flow discourager  248  from hot gas ingestion as well as high temperatures of the mainstream gas flow. Exemplary embodiments may particularly prevent or reduce creep of the aft flow discourager  248  while not adding additional material to the rotor disk  246 , turbine platform  244 , and/or the aft flow discourager  248 . Exemplary embodiments may also maintain the aft flow discourager  248  under centrifugal load. Exemplary embodiments may minimize the amount of air necessary to cool the gas turbine engine  100  and increase efficiency. Additionally, because of the simplicity of the design, the systems and methods disclosed herein can be readily incorporated on new design engines or it can be economically retrofitted on existing engines. The gas turbine engine assemblies produced according to exemplary embodiments may find beneficial use in many industries including aerospace, but also including industrial applications such as electricity generation, naval propulsion, pumping sets for gas and oil transmission, aircraft propulsion, automobile engines, and/or stationary power plants. 
     While at least one exemplary embodiment has been presented in the foregoing detailed description of the invention, it should be appreciated that a vast number of variations exist. It should also be appreciated that the exemplary embodiment or exemplary embodiments are only examples, and are not intended to limit the scope, applicability, or configuration of the invention in any way. Rather, the foregoing detailed description will provide those skilled in the art with a convenient road map for implementing an exemplary embodiment of the invention. It being understood that various changes may be made in the function and arrangement of elements described in an exemplary embodiment without departing from the scope of the invention as set forth in the appended claims.

Technology Classification (CPC): 5