Patent Publication Number: US-7588413-B2

Title: Upstream plasma shielded film cooling

Description:
BACKGROUND OF THE INVENTION 
   1. Field of the Invention 
   The invention relates to film cooling of hot surfaces such as those found in hot aircraft gas turbine engine components and, particularly, to film cooling holes such as those found in combustor liners and turbine nozzle airfoils in gas turbine engines. 
   2. Description of Related Art 
   A typical gas turbine engine of the turbofan type generally includes a forward fan and a booster or low pressure compressor, a middle core engine, and a low pressure turbine which powers the fan and booster or low pressure compressor. The core engine includes a high pressure compressor, a combustor and a high pressure turbine in a serial flow relationship. The high pressure compressor and high pressure turbine of the core engine are connected by a high pressure shaft. High pressure air from the high pressure compressor is mixed with fuel in the combustor and ignited to form a very hot high energy gas stream. The gas stream flows through the high pressure turbine, rotatably driving it and the high pressure shaft which, in turn, rotatably drives the high pressure compressor. 
   The gas stream leaving the high pressure turbine is expanded through a second or low pressure turbine. The low pressure turbine rotatably drives the fan and booster compressor via a low pressure shaft. The low pressure shaft extends through the high pressure rotor. Most of the thrust produced is generated by the fan. Marine or industrial gas turbine engines have low pressure turbines which power generators, ship propellers, pumps and other devices while turboprops engines use low pressure turbines to power propellers usually through a gearbox. 
   The high pressure turbine has a turbine nozzle including at least one row of circumferentially spaced apart airfoils or vanes radially extending between radially inner and outer bands. The vanes are usually hollow having an outer wall that is cooled with cooling air from the compressor. Hot gases flowing over the cooled turbine vane outer wall produces flow and thermal boundary layers along hot outer surfaces of the vane outer wall and end wall hot surfaces of the inner and outer bands over which the hot gases pass. 
   Film cooling is widely used in gas turbine hot components, such as combustor liners, turbine nozzle vanes and bands, turbine blades, turbine shrouds, and exhaust nozzles and exhaust nozzle liners such as those used for afterburning engines. Film cooling is used to inject cooler air through film cooling holes or slots to form an insulating layer on the component hot surface and reduce the direct contact with the hot gases flowing over the component surface. The film cooling holes are typically angled in a downstream direction so that the cooling air is injected into the boundary layer along or as close as possible to the hot surface. The cooling film flow can mix with the hot gas and reduce its effectiveness as it flows in the downstream direction. One method of reducing film mixing with hot gases is to have an aft facing step at the upstream of the holes or slots to shield the film flow. This method has been used in combustor liners where the gas velocity is lower, but not in the turbine airfoils where the gas velocity is higher. The aft facing step is a physical intrusion from the aerodynamic surfaces. In high speed applications, the physical intrusion can cause significant aerodynamic losses. It is desirable to have a device that can provide the similar shielding effect for the film cooling without physical intrusion for maintaining aerodynamic efficiency. 
   SUMMARY OF THE INVENTION 
   An upstream plasma boundary layer shielding system includes film cooling apertures disposed through a wall and angled in a downstream direction from a cold surface of the wall to an outer hot surface of the wall. A plasma generator located upstream of the film cooling apertures is used for producing a plasma extending over the film cooling apertures. 
   In an exemplary embodiment of the system, the plasma generator is mounted on the wall and includes inner and outer electrodes separated by a dielectric material. An AC power supply is connected to the electrodes to supply a high voltage AC potential to the electrodes. The dielectric material being disposed within a groove in the outer hot surface of the wall. 
   A more particular embodiment of the system further includes a gas turbine engine vane including the wall defining at least in part a hollow airfoil of the vane. The airfoil extends radially in a spanwise direction between radially inner and outer bands and in the downstream direction and in a chordwise direction between opposite leading and trailing edges. The airfoil may be part of a high pressure turbine nozzle vane. The plasma generator may be mounted on the airfoil with the dielectric material disposed within a spanwise extending groove in an outer hot surface of the airfoil. 
   Another more particular embodiment of the system further includes the wall being annular and defining at least in part a gas turbine engine combustor liner and the groove being annular. 
   A method for operating an upstream plasma boundary layer shielding system includes energizing a plasma generator to form a plasma extending in a downstream direction over film cooling apertures disposed through a wall and along an outer hot surface of the wall. The plasma generator may be operated in steady state or unsteady modes. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The foregoing aspects and other features of the invention are explained in the following description, taken in connection with the accompanying drawings where: 
       FIG. 1  is a longitudinal, sectional view illustration of exemplary embodiment of an aircraft gas turbine engine with an upstream plasma boundary layer shielding system illustrated for nozzle vanes of a high pressure turbine section of the engine. 
       FIG. 2  is an enlarged view of the nozzle vanes and the upstream plasma boundary layer shielding system illustrated in  FIG. 1 . 
       FIG. 3  is an enlarged perspective view of the vanes and plasma generators of the upstream plasma boundary layer shielding system illustrated in  FIG. 2 . 
       FIG. 4  is a cross sectional view illustration through the vanes illustrated in  FIG. 3 . 
       FIG. 5  is a schematical and partial graphical illustration of the upstream plasma boundary layer shielding system with the plasma generators illustrated in  FIG. 4  energized and a boundary layer. 
       FIG. 6  is a graphical illustration of the boundary layer without the plasma generators. 
       FIG. 7  is a perspective view of a gas turbine engine liner with an upstream plasma boundary layer shielding system. 
   

   DETAILED DESCRIPTION OF THE INVENTION 
   Illustrated in  FIG. 1  is an exemplary turbofan gas turbine engine  10  circumscribed about an engine centerline axis  8  and having a fan  12  which receives ambient air  14 , a booster or low pressure compressor (LPC)  16 , a high pressure compressor (HPC)  18 , a combustor  20  which mixes fuel with the air  14  pressurized by the HPC  18  for generating combustion gases or gas flow  19  which flows downstream through a high pressure turbine (HPT)  22 , and a low pressure turbine (LPT)  24  from which the combustion gases are discharged from the engine  10 . The HPT  22  is joined to the HPC  18  to substantially form a high pressure rotor  29 . A low pressure shaft  28  joins the LPT  24  to both the fan  12  and the low pressure compressor  16 . The second or low pressure shaft  28  which is at least in part rotatably disposed co-axially with and radially inwardly of the first or high pressure rotor. 
   Illustrated in  FIGS. 2 and 3  is the turbine nozzle  30  of the high pressure turbine  22  through which the hot gas flow  19  is discharged into from the combustor  20 . The main combustor  20  includes inner and outer combustor liners  74 ,  76 . The exemplary embodiment of the turbine nozzle  30 , which is also more generally referred to as a vane assembly  31 , illustrated in  FIGS. 2 and 3  includes a row  33  of circumferentially spaced apart vanes  32  extending radially in a spanwise direction S between radially inner and outer bands  38 ,  40 , respectively. In the exemplary embodiment of the turbine nozzle  30  illustrated herein, the bands and vanes are formed in circumferential segments  42  typically, with two vanes  32  per segment  42 . There may be more than two segments and the segments typically have axial split lines suitably joined together by conventional spline seals therebetween. A portion of compressor discharge air  45  is used to supply pressurized cooling air  35  to the turbine nozzle  30  for cooling the various components thereof including the hollow airfoils  39  and inner and outer bands. Cooling air  35  is also used to film cool an annular shroud  72  surrounding rotatable blade tips  82  of the high pressure turbine  22 . 
   Referring to  FIGS. 3 and 4 , each airfoil  39  includes an outer wall  26  having a pressure side  46  and a circumferentially opposite suction side  48  which extend axially in a chordwise direction C between opposite leading and trailing edges LE, TE, respectively. The airfoils  39  and the outer walls  26  extend radially in a spanwise direction S between the inner and outer bands  38 ,  40 . The bands are typically integrally cast with the corresponding vanes during initial manufacture thereof. The hot combustion gas flow  19  pass through flow passages  50  between the airfoils  39 . The flow passages  50  are bound by inboard hot surfaces  52 , with respect to the gas flow  19 , of the inner and outer bands  38 ,  40  and outer hot surfaces  54  of then outer wall  26  along the pressure and suction sides  46 ,  48  of the airfoils  39 . 
   The hot combustion gas flow  19  flowing over the cooled turbine vanes  32  and outer walls  26  form a flow boundary layer  60  along the inboard hot surfaces  52  of the inner and outer bands  38 ,  40  and, as schematically illustrated in  FIG. 6 , along the outer hot surfaces  54  of the pressure and suction sides  46 ,  48  of the outer walls  26 . There is a velocity gradient V and a gas temperature gradient T within the flow boundary layer  60  adjacent to the outer hot surfaces  54  of the pressure and suction sides  46 ,  48  of the outer walls  26 . The gas temperature gradient T and the hot gas flow  19  causes unwanted and undesirable heating along the outer hot surfaces  54  of the pressure and suction sides  46 ,  48  of the outer walls  26 . The gas temperature gradient T results in undesirable heat transfer from the hot gas flow  19  to the relatively colder yet hot outer walls  26 . 
   The outer walls  26  are film cooled by using pressurized cooling air  35  which is a portion of the compressor discharge air  45  from last high pressure compressor stage  43  at a downstream end of the high pressure compressor  18  as illustrated in  FIGS. 1 and 2 . The portion of the compressor discharge air  45  flows around the outer combustor liner  76  and through liner apertures  44  in a downstream flange  47  of the outer combustor liner  76  into a cooling air plenum  56 . The portion of the compressor discharge air  45  that flows into the cooling air plenum  56  is used as the cooling air  35  and flows into hollow interiors  41  of the airfoils  39 . 
   Film cooling apertures  49 , such as cylindrical or other shaped holes or slots, are disposed through the outer wall  26  on the pressure and suction sides  46 ,  48  of the airfoils  39  as illustrated in  FIGS. 2 ,  3 , and  4 . The film cooling apertures  49  are used to flow cooling air  35  across the outer wall  26  and form a thermally protective cooling film  37  over the outer hot surface  54  of the wall  26 . An upstream plasma boundary layer shielding system  11  illustrated in  FIG. 1  is designed to shield the film cooled outer hot surfaces  54  of the walls  26  of the airfoils  39  in the turbine nozzle  30  of the high pressure turbine  22 . The upstream plasma boundary layer shielding system  11  is also designed to shield the film cooled outer hot surfaces  54  of walls  26  such as those found in the combustor  20  as well as other film cooled hot surfaces of other gas turbine engine components and non-gas turbine engine film cooled walls. 
   The film cooling apertures  49  are angled in a downstream direction with respect to the hot gas flow  19 . The film cooling apertures  49  extend across the wall  26  from a cold surface  59  of the wall  26  to the outer hot surface  54  of the wall  26  in a generally downstream direction D. The terms cold surface  59  and outer hot surface  54  are used to designate which of the surfaces are relatively cold and hot during operation of the engine or heating of the wall  26  and does not reflect their relative temperatures when the system  11  is not being operated. The film cooling apertures  49  are typically shallow with respect to the wall  26  and angled in the downstream direction in order to entrain the film cooling air  35  in the boundary layer along the outer hot surface  54  and form the cooling film  37  over the hot surface. An electronic controller  51  may be used to control and turn on and off plasma generators  2  and an active clearance control system if the engine has one. 
   The upstream plasma boundary layer shielding system  11  illustrated in  FIGS. 2-5  is designed to form a virtual aerodynamic shield for the cooling film  37  over the outer hot surface  54  of the wall  26 . The upstream plasma boundary layer shielding system  11  illustrated herein includes plasma generators  2  located on the outer hot surface  54  of the wall  26  upstream U of the film cooling apertures  49  as illustrated in  FIG. 5 . A plasma generator  2  is disposed on each of the suction sides  48  and the pressure sides  46  of the airfoils  39  upstream or forward of the film cooling apertures  49 . The plasma generators  2  produce an airfoil outer surface conforming plasma  90  along each of the outer hot surfaces  54  of the pressure and suction sides  46 ,  48  of the airfoils  39 . The upstream plasma boundary layer shielding system  11  lifts the flow boundary layer  60  off of and up and away from the outer hot surfaces  54  of the outer walls  26  of the airfoils  39  as illustrated in  FIG. 5 . This forms a slip boundary layer  70  for the gas flow  19  to flow over and protects the cooling film  37  and further reduces the amount of heat transferred to the wall  26 . 
   The slip boundary layer  70  provides an interface  68  between the gas flow  19  and the outer hot surface  54  of the outer wall  26  and the interface  68  is not a solid surface when the plasma generators  2  are turned on or energized. The flow boundary layer  60  and its velocity and gas temperature gradients V, T are separated from the outer hot surface  54  by the slip layer  70  when the plasma generators  2  are energized as illustrated in  FIG. 5  whereas the gradients directly contact the outer hot surface  54  when the plasma generators  2  are not energized as illustrated in  FIG. 6 . 
   Shielding of the cooling film  37  on the outer hot surfaces  54  reduces surface heat transfer between the gas flow  19  and the outer hot surfaces  54  of the outer walls  26  of the airfoils  39  due to the flow boundary layer  60 . Reduction of heat transfer improves component life of the vane or other upstream plasma shielded film cooled component and lowers cooling flow requirement for the component and, thus, improves engine efficiency. 
   Referring to  FIG. 5 , an exemplary embodiment of the plasma generator  2  illustrated herein includes the plasma generators  2  mounted on the outer walls  26  of the vanes  32 . Each of the plasma generators  2  includes inner and outer electrodes  3 ,  4  separated by a dielectric material  5 . The dielectric material  5  is disposed within spanwise extending grooves  6  in the outer hot surfaces  54  of the outer walls  26  of the vanes  32 . An AC power supply  100  is connected to the electrodes to supply a high voltage AC potential to the electrodes. 
   When the AC amplitude is large enough, the gas flow  19  ionizes in a region of largest electric potential forming the plasma  90 . The plurality of plasma generators  2  produce an outer surface conforming plasma  90  which covers a substantial portion of the outer hot surface  54  of the vane  32 . The plasma  90  generally begins at an edge  102  of the outer electrode  4  which is exposed to the gas flow  19  and spreads out over an area  104  projected by the outer electrode  4  which is covered by the dielectric material  5 . The plasma  90  in the presence of an electric field gradient produces a force on the gas flow  19  located between the outer hot surface  54  and the plasma  90  inducing a virtual aerodynamic shield for the cooling film over the outer hot surface  54  of the outer wall  26  of the airfoil  39 . The induced aerodynamic shield and resulting change in the pressure distribution forms the slip boundary layer  70  for the gas flow  19  to flow over the cooling film  37 . It is known that airfoils using plasma generators have been shown to prevent flow separation over the airfoils. 
   When the plasma generators  2  are turned on, the velocity gradient V at the interface  68  is smaller than when the plasma generators  2  are off. Similarly, the temperature gradient T at the interface  68  is also smaller when the plasma generators  2  are on than when the plasma generators  2  are off. Therefore, heating from the hot gas flow  19  to the outer hot surfaces  54  of the suction sides  48  of the outer walls  26  of the airfoils  39  will also be smaller when the plasma generators  2  are on than when the plasma generators  2  are off. The plasma generators  2  may be operated in either steady state or unsteady modes. 
   The upstream plasma boundary layer shielding system  11  is illustrated in  FIGS. 1-6  for use with the airfoils  39  of the turbine nozzle  30  of the high pressure turbine  22  and, more particularly, for use on both the pressure and suction sides  46 ,  48  of the airfoil&#39;s outer or hot wall. The upstream plasma boundary layer shielding system  11  may also be used along the inboard hot surfaces  52  of the inner and outer bands  38 ,  40  and on the inner and outer combustor liners  74 ,  76  of the main combustor  20  illustrated in  FIG. 1 . The upstream plasma boundary layer shielding system  11  may also be used on turbine nozzle airfoils in other stages of a high pressure turbine and in an afterburner combustor liner. 
   An afterburner combustor or exhaust nozzle liner is illustrated in U.S. Pat. No. 5,465,572 and main combustor liner is more particularly illustrated in U.S. Pat. No. 5,181,379. A portion  64  of a gas turbine engine liner  66  is exemplified by an annular combustor liner  66  which may be from a main or afterburner combustor liner or an exhaust nozzle liner, as illustrated in  FIG. 7 . Combustor and exhaust nozzle liners are typically annular and circumscribed about the engine centerline axis  8 . The dielectric material  5  is disposed within an annular groove  6  in inwardly facing hot surfaces or the wall that makes up the liners. Film cooling apertures  49 , illustrated as being cylindrical, are disposed through the outer wall  26  which is illustrated as being annular. 
   The plasma generator  2  is located on the outer hot surface  54  of the wall  26  upstream U of the film cooling apertures  49 . The film cooling apertures  49  are angled in a downstream direction with respect to the hot gas flow  19 . The film cooling apertures  49  extend across the wall  26  from a cold surface  59  of the wall  26  to the outer hot surface  54  of the wall  26  in a generally downstream direction D. The film cooling apertures  49  are typically shallow with respect to the wall  26  and angled in the downstream direction in order to entrain the film cooling air  35  in the boundary layer along the outer hot surface  54  and form the cooling film  37  over the hot surface. The cooling air  35  flows through the film cooling apertures  49  in a radially inwardly and downstream direction. The upstream plasma boundary layer shielding system  11  may also be used in a two dimensional or otherwise shaped gas turbine engine nozzle or exhaust liner. 
   The present invention has been described in an illustrative manner. It is to be understood that the terminology which has been used is intended to be in the nature of words of description rather than of limitation. While there have been described herein, what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein and, it is, therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention.