Patent Abstract:
An exemplary embodiment of the invention is directed to a turbine engine having a first, rotatable turbine rotor assembly, a second, stationary nozzle assembly disposed adjacent thereto and a wheel space which is defined between the first, rotatable turbine rotor assembly and the second, stationary nozzle assembly. The wheel space is operable to receive cooling air therein and includes a sealing feature located on the first rotatable turbine rotor assembly that extends axially into the wheel space to terminate adjacent to a sealing land positioned on the second, stationary nozzle assembly. The sealing feature and the sealing land operate to control the release of cooling air from within the wheel space and the sealing land is constructed of shape memory alloy.

Full Description:
BACKGROUND OF THE INVENTION 
       [0001]    The subject matter disclosed herein relates to gas turbine engines and, more particularly, to temperature and performance management therein. 
         [0002]    In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gas that flows downstream through one or more turbine stages. A turbine stage includes a stationary nozzle having stator vanes that guide the combustion gas through a downstream row of turbine rotor blades. The blades extend radially outwardly from a supporting rotor that is powered by extracting energy from the gas. 
         [0003]    A first stage turbine nozzle receives hot combustion gas from the combustor and directs it to the first stage turbine rotor blades for extraction of energy therefrom. A second stage turbine nozzle may be disposed downstream from the first stage turbine rotor blades, and is followed by a row of second stage turbine rotor blades that extract additional energy from the combustion gas. Additional stages of turbine nozzles and turbine rotor blades may be disposed downstream from the second stage turbine rotor blades. 
         [0004]    As energy is extracted from the combustion gas, the temperature of the gas is correspondingly reduced. However, since the gas temperature is relatively high, the turbine stages are typically cooled by a coolant such as compressed air diverted from the compressor through the hollow vane and blade airfoils for cooling various internal components of the turbine. Since the cooling air is diverted from use by the combustor, the amount of extracted cooling air has a direct influence on the overall efficiency of the engine. It is therefore desired to improve the efficiency with which the cooling air is utilized to improve the overall efficiency of the turbine engine. 
         [0005]    The quantity of cooling air required is dependant not only on the temperature of the combustion gas but on the integrity of the various seals which are disposed between rotating and stationary components of the turbine. Thermal expansion and contraction of the rotor and blades may vary from the thermal expansion of the stationary nozzles and the turbine housing thereby challenging the integrity of the seals. In some cases the seals may be compromised causing excess cooling air to pass into the turbine mainstream gas flow resulting in excess diversion of compressor air translating directly to lower than desired turbine efficiency. 
         [0006]    It is therefore desired to provide a gas turbine engine having improved sealing of gas turbine stationary to rotating component interfaces. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0007]    In an exemplary embodiment of the invention a turbine engine comprises a first, rotatable turbine rotor assembly, a second, stationary nozzle assembly disposed adjacent thereto and a wheel space which is defined between the first, rotatable turbine rotor assembly and the second, stationary nozzle assembly. The wheel space is configured to receive cooling air therein and includes a sealing feature located on the first rotatable turbine rotor assembly that extends axially into the wheel space to terminate adjacent to a sealing land positioned on the second, stationary nozzle assembly. The sealing feature and the sealing land operate to control the release of cooling air from within the wheel space and the sealing land is constructed of shape memory alloy. 
         [0008]    In another embodiment of the invention a turbine engine comprises a first, rotatable turbine rotor assembly, a second, stationary nozzle assembly disposed adjacent thereto and a wheel space defined between the first, rotatable turbine rotor assembly and the second, stationary nozzle assembly and configured to receive cooling air therein. A sealing feature located on the first, rotatable turbine rotor assembly extends axially into the wheel space to terminate adjacent to a sealing land positioned on the second, stationary nozzle assembly. The sealing feature and the sealing land operate to control the release of the cooling air from within the wheel space; the sealing land constructed of shape memory alloy. 
         [0009]    In another embodiment, a turbine engine comprises a turbine housing having an upstream and a downstream end. A stationary nozzle assembly is disposed within the housing in fixed relationship thereto. A turbine rotor assembly is supported within the housing for rotation therein and is operable, during operation of the turbine engine, to thermally expand in the downstream direction relative to the stationary nozzle assembly. A wheel space, defined between the stationary nozzle assembly and the rotatable turbine rotor assembly, is configured to receive cooling air therein. A sealing feature, located on the rotatable turbine rotor assembly and extending axially into the wheel space terminates adjacent to a sealing land positioned on the second, stationary nozzle assembly. The sealing feature and the sealing land operate to control the release of the cooling air from within the wheel space. The sealing land is constructed of shape memory alloy having a composition such that a phase changes from a cold, martensitic state to a hot, austenitic state is within the heat transient of the gas turbine engine. The shape memory alloy is configured as a two-way alloy having a first configuration in the cold, martensitic state and a second configuration in the hot, austenitic state and is operable to maintain the sealing feature adjacent the sealing land during thermal expansion of the turbine rotor assembly. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0010]    The invention, in accordance with preferred and exemplary embodiments, together with further advantages thereof, is more particularly described in the following detailed description taken in conjunction with the accompanying drawings in which: 
           [0011]      FIG. 1  is an axial sectional view through a portion of an exemplary gas turbine engine in accordance with an embodiment of the invention; 
           [0012]      FIG. 2  is an enlarged sectional view through a portion of the gas turbine engine of  FIG. 1 ; 
           [0013]      FIG. 3  is an enlarged sectional view through a portion of the gas turbine engine of  FIG. 1  in a cold, non-operational state; and 
           [0014]      FIG. 4  is an enlarged sectional view through a portion of the gas turbine engine of  FIG. 1  in a hot, operational state. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0015]    Illustrated in  FIGS. 1 and 2  is a portion of a gas turbine engine  10 . The engine is axisymmetrical about a longitudinal, or axial centerline axis and includes, in serial flow communication, a multistage axial compressor  12 , a combustor  14 , and a multi-stage turbine  16 . 
         [0016]    During operation, compressed air  18  from the compressor  12  flows to the combustor  14  that operates to combust fuel with the compressed air for generating hot combustion gas  20 . The hot combustion gas  20  flows downstream through the multi-stage turbine  16 , which extracts energy therefrom. 
         [0017]    As shown in  FIGS. 1 and 2 , an example of a multi-stage axial turbine  16  may be configured in three stages having six rows of airfoils  22 ,  24 ,  26 ,  28 ,  30 ,  32  disposed axially, in direct sequence with each other, for channeling the hot combustion gas  20  therethrough and, for extracting energy therefrom. 
         [0018]    The airfoils  22  are configured as first stage nozzle vane airfoils. The airfoils are circumferentially spaced apart from each other and extend radially between inner and outer vane sidewalls  34 ,  36  to define first stage nozzle assembly  38 . The nozzle assembly  38  is stationary within the turbine housing  40  and operates to receive and direct the hot combustion gas  20  from the combustor  14 . Airfoils  24  extend radially outwardly from the perimeter of a first supporting disk  42  to terminate adjacent first stage shroud  44 . The airfoils  24  and the supporting disk  42  define the first stage turbine rotor assembly  46  that receives the hot combustion gas  20  from the first stage nozzle assembly  38  to rotate the first stage turbine rotor assembly  46 , thereby extracting energy from the hot combustion gas. 
         [0019]    The airfoils  26  are configured as second stage nozzle vane airfoils. The airfoils are circumferentially spaced apart from each other and extend radially between inner and outer vane sidewalls  48  and  50  to define second stage nozzle assembly  52 . The second stage nozzle assembly  52  is stationary within the turbine housing  40  and operates to receive the hot combustion gas  20  from the first stage turbine rotor assembly  46 . Airfoils  28  extend radially outwardly from a second supporting disk  54  to terminate adjacent second stage shroud  56 . The airfoils  28  and the supporting disk  54  define the second stage turbine rotor assembly  58  for directly receiving hot combustion gas  20  from the second stage nozzle assembly  52  for additionally extracting energy therefrom. 
         [0020]    Similarly, the airfoils  30  are configured as third stage nozzle vane airfoils circumferentially spaced apart from each other and extending radially between inner and outer vane sidewalls  60  and  62  to define a third stage nozzle assembly  64 . The third stage nozzle assembly  64  is stationary within the turbine housing  40  and operates to receive the hot combustion gas  20  from the second stage turbine rotor assembly  58 . Airfoils  32  extend radially outwardly from a third supporting disk  66  to terminate adjacent third stage shroud  68 . The airfoils  32  and the supporting disk  66  define the third stage turbine rotor assembly  70  for directly receiving hot combustion gas  20  from the third stage nozzle assembly  64  for additionally extracting energy therefrom. The number of stages utilized in a multistage turbine  16  may vary depending upon the particular application of the gas turbine engine  10 . 
         [0021]    As indicated, first, second and third stage nozzle assemblies  38 ,  52  and  64  are stationary relative to the turbine housing  40  while the turbine rotor assemblies  46 ,  58  and  70  are mounted for rotation therein. As such, there are defined between the stationary and rotational components, cavities that may be referred to as wheel spaces. Exemplary wheel spaces  72  and  74 , illustrated in  FIG. 2 , reside on either side of the second stage nozzle assembly  52  between the nozzle assembly and the first stage turbine rotor assembly  46  and the nozzle assembly and the second stage rotor assembly  58 . 
         [0022]    The turbine airfoils as well as the wheel spaces  72 ,  74  are exposed to the hot combustion gas  20  during operation of the turbine engine  10 . To assure desired durability of such internal components they are typically cooled. For example, second stage nozzle airfoils  26  are hollow with walls  76  defining a coolant passage  78 . In an exemplary embodiment, a portion of compressed air from the multistage axial compressor  12  is diverted from the combustor and used as cooling air  80 , which is channeled through the airfoil  26  for internal cooling. Extending radially inward of the second stage inner vane sidewall  48  is a diaphragm assembly  82 . The diaphragm assembly includes radially extending side portions  84  and  86  with an inner radial end  87  closely adjacent the rotor surface  88 . An inner cooling passage  90  receives a portion of the cooling air  80  passing through the airfoil coolant passage  78  and disperses the cooling air into the wheel spaces  72  and  74  to maintain acceptable temperature levels therein. Sealing features  92  and  94 , referred to as “angel wings”, are disposed on the upstream and downstream sides of the first stage turbine airfoils  24 . Similarly, sealing features  96  and  98  are disposed on the upstream and downstream sides of the second stage turbine airfoils  28 . The sealing features, or angel wings, extend in an axial direction and terminate within their associated wheel spaces closely adjacent to complementary sealing lands such as  100  and  102 , mounted in and extending from radially extending side portions  84 ,  86  of the second stage diaphragm assembly  82 . During operation of the turbine engine, leakage of cooling air  80 , flowing into the wheel spaces  72  and  74  from the inner cooling passage  90  of the diaphragm assembly  82 , is controlled by the close proximity of the upstream and downstream sealing features  96 ,  94  and the sealing lands  100 ,  102 . Similar sealing features and sealing lands may also be used between stationary and rotating portions of the other turbine stages of the turbine engine  10 . 
         [0023]    During operation of the gas turbine engine  10 , especially as the temperature of the engine transitions from a cold state to a hot state following start-up, the various components of the engine, already described above, may experience some degree of thermal expansion resulting in dimensional changes in the engine  10  which must be accounted for. For instance, as the temperature rises, the entire turbine rotor assembly  104  may expand axially relative to the fixed nozzle assemblies as well as the turbine housing  40 . Due to the manner in which the turbine rotor assembly  104  is supported within the turbine housing  40 , such axial expansion is primarily in the down stream direction relative to the housing,  FIG. 1 . As a result of the downstream relative movement, the axial over-lap spacing between the downstream sealing features  94  of first stage turbine rotor assembly  46  and the second stage upstream sealing land  100  may increase, resulting in a decrease in the leakage of cooling air  80  into the main gas stream  20  from wheel space  72 . Conversely, the axial over-lap spacing between the second stage downstream sealing land  102  and the upstream sealing feature  96  of the second stage turbine rotor assembly  58  may decrease. Baring contact, the increase/decrease between sealing features is of minor consequence. However, since the cooling air  80  is diverted air from the axial compressor, its usage for purposes other than combustion will directly influence the efficiency of the gas turbine engine  10  and the designed operation of the wheel spaces. Each wheel space is designed to maintain a specific flow of cooling air to prevent the ingestion of the main gas stream  20  into the wheel space. Therefore, the decrease in axial over-lap spacing between the upstream sealing features  96  of second stage turbine rotor assembly  58  and the second stage downstream sealing land  102  is undesirable because the incorrect amount of flow is delivered to this wheel space  74 . Accordingly, wheel space  74  with its decrease in axial over-lap distance will leak more than the designed flow into the main gas stream  20 . 
         [0024]    In one exemplary embodiment, the second stage downstream sealing land  102  comprises a band that is constructed of a two-way shape memory metal such as a nickel-titanium (“NiTi”) alloy. Shape memory alloy can exist in two different, temperature dependant crystal structures or phases (i.e. martensite (lower temperature) and austenite (higher temperature)), with the temperature at which the phase change occurs dependant upon the composition of the alloy. Two-way shape memory alloy has the ability to recover a preset shape upon heating above the transformation temperature and to return to a certain alternate shape upon cooling below the transformation temperature. Sealing land  102  is configured using a NiTi alloy having a phase change within the heat transient of the gas turbine engine  10 . Through a process of mechanical working and heat treatment, the land  102  is subject to a programming process in which the martensite configuration has an axially shorter length than the austenite configuration, which is axially longer. In some cases the martensite configuration may also be programmed to have a radially differing position relative to the radial sealing feature  96  than in the austenite configuration. As the gas turbine engine  10  transitions from cold to hot following start up, the sealing land  102  will proceed through its martensitic phase  FIG. 3 , to its austenitic phase  FIG. 4 , resulting in axial growth of the land and maintenance of the close physical spacing between the upstream sealing features  96  of second stage turbine rotor assembly  58  and the second stage downstream sealing land  102  regardless of the downstream axial growth of the turbine rotor assembly  104 . The result is reduced passage of cooling air  80  from within the downstream wheel space  74  between second stage turbine rotor assembly  58  and the diaphragm assembly  82  of the second stage nozzle assembly  52 , thereby improving the efficiency of the gas turbine engine and maintaining control of the wheel space cooling air flows. It is contemplated that, if desirable, the sealing land  102  may also be designed to include a radial as well as an axial change in clearance as the gas turbine engine  10  transitions from cold to hot. 
         [0025]    In another embodiment of the invention, the second stage downstream sealing land  102  comprises a band that is constructed of a one-way shape memory metal such as a nickel-titanium (“NiTi”) alloy. Like two-way shape memory alloy, one-way shape memory alloy can exist in two different, temperature dependant crystal structures or phases (i.e. martensite (lower temperature) and austenite (higher temperature), with the temperature at which the phase change occurs dependant upon the composition of the alloy. Unlike two way shape memory alloy, one way allow has the ability to recover a preset shape upon heating above the transformation temperature following its mechanical deformation in the cold, martensite state. Upon cooling, the result of the mechanical deformation is erased. Sealing land  102  is configured using a NiTi alloy having a phase change within the heat transient of the gas turbine engine  10 . As the gas turbine engine  10  transitions from hot to cold following shutdown, the sealing land  102  will transition from its austenitic to its martensite state. Cooling of the turbine rotor assembly  104  results in the axial over-lap spacing between the sealing lands  102  and upstream sealing features  96  of second stage turbine rotor assembly  58  to increase. Following transition to the cold, martensitic phase the sealing land  102  may contact the sealing features  96  resulting in deformation of the sealing land. Following re-start of the gas turbine engine  10  and passage of the sealing land  102  through its martensitic to austenitic phase change the second stage downstream sealing land  102  will return to its un-deformed, initial state in close physical proximity to the upstream sealing features  96  of second stage turbine rotor assembly  58 . The result is reduced leakage of cooling air  80  from within the downstream wheel space  74  between second stage turbine rotor assembly  58  and the diaphragm assembly  82  of the second stage nozzle assembly  52 , thereby improving the efficiency of the gas turbine engine and maintaining control of the wheel space cooling air flows. 
         [0026]    While exemplary embodiments of the invention have been described with application primarily to a second stage of a multi-stage turbine, the focused description is for simplification only and the scope of the invention is not intended to be limited to that single application. The application of the described invention can be applied to similar turbine engine assemblies and components throughout the various stages. 
         [0027]    While exemplary embodiments of the invention have been described with reference to shape memory alloys of a nickel-titanium composition, other compositions such as nickel-metallic cobalt, copper-zinc or others, which exhibit suitable behavior at the desired temperatures of the turbine engine, may be utilized. In addition, the above description has been made with reference to an axial growth component in the seal land. It is recognized that due to the versatility of the shape memory alloys, the sealing land  102  may include a radial as well as an axial change in clearance from cold to hot. 
         [0028]    This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.

Technology Classification (CPC): 5