Variable pitch turbofan engine and a method for operating same

• US 3,946,554 A
• Filed: 09/06/1974
• Issued: 03/30/1976
• Est. Priority Date: 09/06/1974
• Status: Expired due to Term

• Current Assignee: General Electric Company
• Original Assignee: General Electric Company
• Inventors: Neumann, Gerhard
• Primary Examiner(s): Freeh, William L.
• Assistant Examiner(s): Garrett, Robert E.
• Application Number:
  US05/503,752
• Time in Patent Office: 571 Days
• Field of Search: 60/226 R, 60/226 A, 60/262, 60/204, 415/1, 415/149 A, 415/149 R, 415/130
• US Class Current: 60/204
• CPC Class Codes: F01D 17/162 for axial flow, i.e. the vanes turning around axes which are essentially perpendicular to the rotor centre line F01D17/167 takes precedence
• CPC Class Codes: F01D 7/00 Rotors with blades adjustable in operation; Control thereof for reversing F01D1/30
• CPC Class Codes: F02C 9/20 by throttling; by adjusting vanes
• CPC Class Codes: F02K 1/66 using reversing fan blades
• CPC Class Codes: F02K 3/04 the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
• CPC Class Codes: F04D 27/0246 by varying geometry within the pumps, e.g. by adjusting vanes
• CPC Class Codes: F04D 29/323 adjustable
• CPC Class Codes: F04D 29/563 specially adapted for elastic fluid pumps
• CPC Class Codes: F05D 2220/36 specially adapted for the fan of turbofan engines
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

Variable geometry guide vanes are provided in a variable pitch turbofan assembly. A method is disclosed for reducing aerodynamically induced stresses in the variable pitch fan blades as the blades are changed between the forward and reverse pitch modes. The guide vane geometry is scheduled to change in a manner to control the flow into and out of the fan blades in order to establish a favorable aerodynamic environment during the transition between the forward and the reverse pitch modes.

Claims

1. In a method of operating a gas turbine engine having a rotatable fan and in which at least one stage of fan blades is rotating to provide a forward propulsive thrust, the pitch of the rotating stage of fan blades being changed for the purpose of reversing the direction of gas flow therethrough, the step of:

reversing the pitch of at least one stage of essentially radial, nonrotatable, variable guide vanes, which are axially disposed with respect to said rotating fan blades, in a predetermined relationship with the change in pitch of said fan blades to reduce aerodynamically induced stresses in said fan blades as said fan blades change between the forward and the reverse thrust modes.

2. The improved method of claim 1 additionally comprising the step of reversing the pitch of two stages of guide vanes, one disposed axially upstream and the other disposed axially downstream of the fan blades in the forward thrust mode, wherein said scheduling is accomplished in a predetermined relationship with the change in pitch of said fan blades to further reduce aerodynamically induced stresses as said fan blades change between the forward and the reverse thrust modes.

3. The improved method of claim 1 further characterized in that the change in pitch of said guide vanes is achieved by rotating said guide vanes about their essentially longitudinal axes.

4. The improved method of claim 1 further characterized in that the change in pitch of said guide vanes is achieved by changing the camber of said guide vanes.

5. The improved method of claim 3 still further characterized in that said fan blades and said variable guide vanes are scheduled to change pitch in the same direction of rotation about their blade longitudinal axes.

6. In a gas turbine engine incorporating a rotatable fan with blades operating at a first pitch angle during some portions of operation to provide forward thrust and a second pitch angle during other portions of operation to provide reverse thrust, the improvement comprising:

at least one stage of essentially radial, nonrotatable guide vanes axially disposed with respect to said fan, said vanes being adjustable in geometry about the vane longitudinal axes, and

control means to reverse the pitch of said vanes in a predetermined relationship with the change in pitch of the fan blades as the fan is changed between the forward and reverse thrust modes, so as to reduce aerodynamically induced stresses during the transition between the forward and reverse thrust modes.

7. The gas turbine engine of claim 6 wherein said guide vanes are axially disposed upstream of the fan blades in the forward thrust mode.

8. The gas turbine engine of claim 6 wherein said guide vanes are axially disposed downstream of the fan blades in the forward thrust mode.

9. The gas turbine engine of claim 6 additionally comprising two stages of essentially radial, nonrotatable guide vanes, one disposed axially upstream and the other disposed axially downstream with respect to said fan blades, said second stage being variable in pitch and operatively connected to said control means such that variation in pitch cooperates to further reduce stresses as the fan blades are changed between the forward and reverse thrust modes.

10. The gas turbine engine of claim 9 wherein the change in pitch of said vanes is achieved by a rotation of said guide vanes about their essentially longitudinal axes.

11. The gas turbine engine of claim 9 wherein the change in pitch of said vanes is achieved by a change in camber of said guide vanes.

12. A turbofan engine comprising:

a core engine for use in a nacelle, said core engine including, in series flow relationship, an axial flow compressor having rotor and stator assemblies, a combustor to which pressurized air from the compressor and fuel are delivered for combustion therein to provide a hot gas stream, a first turbine downstream of said combustor adapted to be rotatably driven by said hot gas stream, and a shaft drivingly connecting said compressor rotor and said first turbine;

a second turbine rotatable independently of the rotatable elements of the core engine, said second turbine also for use in said nacelle, disposed downstream of said first turbine and adapted to be rotatably driven by said hot gas stream;

a generally annular duct having an inlet and an outlet and for disposition about at least a portion of said core engine;

a fan assembly disposed in said duct for pressurizing air flow through said duct, said second turbine drivingly connected to said fan, and including at least one stage of radially extending fan blades adapted to be variable in pitch, said fan blades operating at a first pitch angle during certain portions of operation to provide forward thrust and at a second pitch angle during other portions of operation to provide reverse thrust;

at least one stage of essentially radial, nonrotatable inlet guide vanes axially disposed with respect to said fan blades, said stage of guide vanes being adjustable in pitch about their respective blade longitudinal axes; and

control means to reverse the pitch of said stage of vanes in a predetermined relationship with the change in pitch of the fan as the fan is changed between the forward and reverse thrust modes, to reduce aerodynamically induced stresses during the transition.

13. The gas turbine engine of claim 12 wherein said guide vanes are axially disposed upstream of the fan blades in the forward thrust mode.

14. The gas turbine engine of claim 12 wherein said guide vanes are axially disposed downstream of the fan blades in the forward thrust mode.

15. The gas turbine engine of claim 12 additionally comprising two stages of essentially radial, nonrotatable guide vanes, one disposed axially upstream and the other disposed axially downstream with respect to said fan blades, said second stage being variable in pitch and operatively connected to said control means such that variation in pitch cooperates to further reduce stresses as the fan blades are changed between the forward and reverse thrust modes.

16. A method of operating a gas turbine engine which includes the steps of:

pressurizing air flowing through a fan assembly disposed in a generally annular duct having an inlet and an outlet, said assembly comprising a stage of variable pitch fan blades and at least one stage of essentially radial, nonrotatable, variable pitch guide vanes axially disposed with respect to said fan blades,

further pressurizing a first predetermined portion of said air flow through an axial flow compressor having rotor and stator assemblies,

mixing the pressurized air from the compressor with a vaporized fuel to form a combustible mixture,

combusting said mixture in a combustor to provide a hot gas stream,

expanding said hot gas stream through a first turbine disposed downstream of said combustor, adapted to be rotatably driven by said hot gas stream, and drivingly connected through a first shaft to said compressor rotor,

further expanding said hot gas stream through a second turbine disposed downstream of said first turbine, independently rotatable from the compressor rotor, and drivingly connected through a second shaft to said fan assembly,

still further expanding said hot gas stream through a first nozzle disposed downstream of said second turbine, and

expanding a second predetermined portion of said pressurized air flow from said fan assembly through a second annular nozzle disposed downstream of said fan assembly,

changing the pitch of said fan blades for the purpose of reversing the direction of gas flow therethrough, and

reversing the pitch of said guide vanes in a predetermined relationship with the change in pitch of said fan blades to reduce aerodynamically induced stresses in said fan blades as said fan blades change between the forward and the reverse thrust modes.

17. In a method of operating a gas turbine engine having a rotatable fan and in which at least one stage of the fan blades is rotating to provide a forward propulsive thrust, the pitch of the rotating stage of the fan blades being changed for the purpose of reversing the direction of gas flow therethrough, the step of:

scheduling the change in geometry of at least one stage of essentially radial, nonrotatable, variable guide vanes, which are axially disposed with respect to said rotating fan blades, in a predetermined relationship with the change in pitch of said fan blades to reduce aerodynamically induced stresses in said fan blades as said fan blades change between the forward and the reverse thrust modes, wherein said variable guide vanes are scheduled to be essentially closed when the fan blades are scheduled through their essentially maximum stress orientation.

18. In a method of operating a gas turbine engine having a rotatable fan and in which at least one stage of the fan blades is rotating to provide a forward propulsive thrust, the pitch of the rotating stage of the fan blades being changed for the purpose of reversing the direction of gas flow therethrough, the step of:

scheduling the change in geometry of at least one stage of essentially radial, nonrotatable, variable guide vanes, which are axially disposed with respect to said rotating fan blades, in a predetermined relationship with the change in pitch of said fan blades to reduce the aerodynamically induced stresses in said fan blades as said fan blades change between the forward and reverse thrust modes, wherein said variable guide vanes are scheduled to be essentially closed prior to scheduling the fan blades to their essentially maximum stress orientation.

19. In a gas turbine engine incorporating a rotatable fan with blades operating at a first pitch angle during some portions of operation to provide forward thrust and a second pitch angle during other portions of operation to provide reverse thrust, the improvement comprising:

at least one stage of essentially radial, nonrotatable guide vanes axially disposed with respect to said fan, said vanes being adjustable in geometry about the vane longitudinal axes, and

control means to schedule the change in geometry of said vanes in a predetermined relationship with the change in pitch of the fan blades as the fan is changed between the forward and reverse thrust modes, so as to reduce aerodynamically induced stresses during the transition between the forward and reverse thrust modes, wherein the guide vanes are operatively connected to the fan blades through said control means such that said guide vanes are essentially closed when said fan blades are scheduled through their maximum stress orientation.