Gas turbine engine with supersonic compressor

Abstract

A gas turbine engine having a compressor section using blades on a rotor to deliver a gas at supersonic conditions to a stator. The stator includes one or more of aerodynamic ducts that have converging and diverging portions for deceleration of the gas to subsonic conditions and to deliver a high pressure gas to combustors. The aerodynamic ducts include structures for changing the effective contraction ratio to enable starting even when designed for high pressure ratios, and structures for boundary layer control. In an embodiment, aerodynamic ducts are provided having an aspect ratio of two to one (2:1) or more, when viewed in cross-section orthogonal to flow direction at an entrance to the aerodynamic duct.

1 . A gas turbine engine, comprising:
 (a) a compressor section, said compressor section comprising:
 (1) an inlet for supply of a selected oxidant-containing gas; 
 (2) a rotor extending along a longitudinal axis, the rotor comprising a plurality of impulse blades configured to act on said selected oxidant-containing gas to impart axial and tangential velocity thereto to provide a supersonic gas flow; 
 (3) a stator, said stator comprising
 (i) one or more aerodynamic ducts each having a radially converging portion and a radially diverging portion that, with input of said supersonic gas flow, generate a plurality of shock waves as said oxidant-containing gas passes through said one or more aerodynamic ducts, each of said one or more aerodynamic ducts sized and shaped to decelerate said supersonic gas flow to subsonic conditions, 
 (ii) bypass gas passageway(s) or a geometrically adjustable portion, or both, operable to adjust an effective contraction ratio of some or all of said one or more aerodynamic ducts, and 
 (iii) boundary layer control structures; 
 
   (b) a gas turbine section, said gas turbine section comprising
 (1) a high pressure combustion chamber for receiving said selected oxidant containing gas from said compressor section, 
 (2) a combustor for receiving fuel from a fuel supply and said selected oxidant containing gas from said compressor section and to burn said fuel to create hot pressurized exhaust gases exiting therefrom, and 
 (3) one or more gas turbines operatively affixed to a shaft and configured to receive said hot pressurized exhaust gases and to expand said hot pressurized exhaust gases therethrough, to produce shaft power. 
   
     
     
         2 . The gas turbine engine as set forth in  claim 1 , wherein said bypass gas passageways are positionable between an open, startup condition wherein discharge gas is passed therethrough, and a closed, operating condition which minimizes or stops passage of gas outward through said bypass gas passageways. 
     
     
         3 . The gas turbine engine as set forth in  claim 2 , wherein said bypass gas passageways are fluidly connected to the atmosphere, for discharge thereto. 
     
     
         4 . The gas turbine engine as set forth in  claim 2 , wherein said bypass gas passageways are fluidly connected with one or more return passageways that return said discharge gas directly or indirectly to said gas flow passage. 
     
     
         5 . The gas turbine engine as set forth in  claim 2 , wherein said bypass gas passageways comprise internal bypass gas passageways, wherein said internal bypass gas passageways are fluidly connected internally within or adjacent said one or more aerodynamic ducts to return said discharged gas to said one or more aerodynamic ducts. 
     
     
         6 . The gas turbine engine as set forth in  claim 1 , wherein said bypass gas passageways comprise external passageways fluidly connected with said one or more aerodynamic ducts. 
     
     
         7 . The gas turbine engine as set forth in  claim 1 , wherein said geometrically adjustable portion is positionable between an open, startup condition wherein said radially converging portion of each of said one or more aerodynamic ducts allows sufficient flow of said selected oxidant containing gas through said one or more aerodynamic ducts to establish and position a normal shock within said one or more aerodynamic ducts, and a closed, operating condition in which said radially converging portion of each of said one or more aerodynamic ducts is set to a selected operating position. 
     
     
         8 . The gas turbine engine as set forth in  claim 1 , wherein said geometrically adjustable portion disposed in each of said one or more aerodynamic ducts, by change in position, changes the effective contraction ratio of the aerodynamic duct in which it is disposed. 
     
     
         9 . The gas turbine engine as set forth in  claim 8 , wherein said geometrically adjustable portion further comprises a pivotable member and an actuator, said pivotable member driven by said actuator, and wherein said geometrically adjustable portion is sized and shaped to change the shape of said radially converging portion of the one or more aerodynamic ducts in which the geometrically adjustable portion is disposed when said geometrically adjustable portion is moved with said actuators. 
     
     
         10 . The gas turbine engine as set forth in  claim 1 , further comprising outlet bleed ports in said one or more aerodynamic ducts, and bleed sub-chambers adjacent said one or more aerodynamic ducts, said bleed sub-chambers in fluid communication with said outlet bleed ports, said bleed sub-chambers configured for passage therethrough of gas removed through said outlet bleed ports. 
     
     
         11 . A gas turbine engine, comprising:
 (a) a compressor section, said compressor section comprising:
 (1) an inlet for supply of a selected oxidant containing gas; 
 (2) a rotor extending along a longitudinal axis, the rotor comprising a plurality of impulse blades configured to act on said selected oxidant-containing gas to impart axial and tangential velocity thereto to provide a supersonic gas flow; 
 (3) a stator, said stator comprising
 (i) one or more aerodynamic ducts having a radially converging portion and a radially diverging portion that with input of said supersonic gas flow generate a plurality of shock waves as said oxidant-containing gas passes through said one or more aerodynamic ducts, said one or more aerodynamic ducts sized and shaped to decelerate said supersonic gas flow to subsonic conditions, 
 (ii) means for adjusting the effective contraction ratio of some or all of said one or more aerodynamic ducts, and 
 (iii) means for controlling a boundary layer of gas flowing through some or all of said one or more aerodynamic ducts; 
 
   (b) a gas turbine section, said gas turbine section comprising
 (1) a combustion chamber for receiving said selected oxidant-containing gas from said compressor section, 
 (2) a combustor for receiving fuel from a fuel supply and said selected oxidant containing gas from said compressor section and burn said fuel to create hot pressurized exhaust gases exiting therefrom, and 
 (3) one or more gas turbines operatively affixed to a shaft and configured to receive and expand therethrough said hot pressurized exhaust gases, to produce shaft power. 
   
     
     
         12 . The gas turbine engine as set forth in  claim 11 , wherein said rotor further comprises a shroud for said plurality of impulse blades. 
     
     
         13 . The gas turbine engine as set forth in  claim 11 , wherein said rotor is effectively sealed with said stator, so as to minimize gas leakage during flow therebetween. 
     
     
         14 . The gas turbine engine as set forth in  claim 11 , wherein said gas passing through said plurality of impulse blades is turned by an angle alpha (α) of at least ninety (90) degrees. 
     
     
         15 . The gas turbine engine as set forth in  claim 11 , wherein each of said plurality of impulse blades has a hub end, a tip end, and a trailing edge, and said supersonic gas flow is provided at said trailing edge of each of said plurality of impulse blades from said hub end to said tip end. 
     
     
         16 . Apparatus for a gas turbine engine, comprising:
 a compressor section, said compressor section comprising
 a casing comprising a low pressure gas inlet and a high pressure gas exit; 
 a rotor comprising a plurality of blades and configured to act on a selected oxidant containing gas to impart axial and tangential velocity thereto to provide a supersonic gas flow; and 
 a stator comprising one or more aerodynamic ducts configured for diffusing the supersonic gas flow received therein, said one or more aerodynamic ducts each having a radially converging portion, a radially diverging portion, and an effective contraction ratio, such that, with input of the supersonic gas flow, each of said one or more aerodynamic ducts generates a plurality of oblique shock waves S 1  to S X  and a normal shock wave S N  in said selected oxidant containing gas as said selected oxidant containing gas passes therethrough, said one or more aerodynamic ducts having an inlet relative Mach number for operation associated with a design operating point selected within a design operating envelope for a selected gas composition, gas quantity, and gas compression ratio, said one or more aerodynamic ducts comprising
 (a) bypass gas passageways or geometrically adjustable portions, or both, operable to adjust said effective contraction ratio, and 
 (b) boundary layer control structures comprising one or more of (1) outlet bleed ports for boundary layer removal, (2) inlet jets for energizing a boundary layer by gas injection, and (3) one or more vortex generators. 
 
   
     
     
         17 . The apparatus as set forth in  claim 16 , wherein one or more of said one or more aerodynamic ducts are wrapped helically about said longitudinal axis. 
     
     
         18 . The apparatus as set forth in  claim 16 , wherein each of said one or more aerodynamic ducts comprises a leading edge associated therewith. 
     
     
         19 . The apparatus as set forth in  claim 18 , further comprising:
 a partition wall downstream from said leading edge,   wherein the one or more aerodynamic ducts comprises a plurality of aerodynamic ducts, and said partition wall divides adjacent aerodynamic ducts, and   wherein said leading edge comprises an upstream terminus of said partition wall.   
     
     
         20 . The apparatus as set forth in  claim 16 , further comprising a gas turbine section, said gas turbine section comprising:
 (1) a high pressure combustion chamber for receiving said selected oxidant containing gas from said compressor section,   (2) a combustor for receiving fuel from a fuel supply, said selected oxidant containing gas from said compressor section and to burn said fuel to create hot pressurized exhaust gases exiting therefrom,   (3) one or more gas turbines operatively affixed to a shaft and configured to receive said hot pressurized exhaust gases and to expand said hot pressurized exhaust gases therethrough, to produce shaft power.