Abstract:
A fan assembly for a gas turbine engine includes a turbine rotor adapted to be disposed aft of a core of the gas turbine engine; a row of turbine blades carried by the rotor, each turbine blade extending from the rotor to a tip, the turbine blades adapted to extract energy from a stream of pressurized combustion gases generated by the core; and at least two rows of axially-spaced apart, radially-extending fan blades carried by the row of turbine blades for rotation therewith.

Description:
BACKGROUND OF THE INVENTION  
       [0001]     This invention relates generally to gas turbine engines and more particularly to an aft fan for a gas turbine engine.  
         [0002]     A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot combustion gases. These gases flow downstream to one or more turbines that extract energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. In a turbofan engine, which typically includes a fan placed at the front of the core engine, a high pressure turbine powers the compressor of the core engine. A low pressure turbine is disposed downstream from the high pressure turbine for powering the fan.  
         [0003]     Some prior art engine configurations incorporate an aft fan stage integral with a turbine rotor. There are several potential advantages for this “fan on turbine” configuration, which eliminates the drive shaft required in a front-fan engine. It is desired to have such a fan-on-turbine with a design pressure ratio (i.e. the ratio of total pressure at the fan exit to the total pressure at the fan inlet) of about 2.5 or greater. Unfortunately, the high tip speed required for a single fan stage to produce this pressure ratio is contrary the AN 2  and radius ratio constraints dictated by accepted turbine design practice.  
         [0004]     Accordingly, there is a need for a fan-on-turbine configuration which achieves a high pressure ratio.  
       BRIEF SUMMARY OF THE INVENTION  
       [0005]     The above-mentioned need is met by the present invention, which according to one aspect provides a fan assembly for a gas turbine engine, including: a turbine rotor adapted to be disposed aft of a core of the gas turbine engine; a row of turbine blades carried by the rotor, each turbine blade extending from the rotor to a tip, the turbine blades adapted to extract energy from a stream of pressurized combustion gases generated by the core; and at least two rows of axially-spaced apart, radially-extending fan blades carried by the row of turbine blades for rotation therewith.  
         [0006]     According to another aspect of the invention, a gas turbine engine includes a core for generating a stream of pressurized combustion gases, including in sequential flow order: a compressor, a combustor, and a high-pressure turbine; and a fan assembly having a turbine rotor and disposed aft of the core; a row of turbine blades carried by said rotor, each turbine blade extending from the rotor to a tip, said turbine blades adapted to extract energy from the combustion gases; and at least two rows of axially-spaced apart, radially-extending fan blades carried by the row of turbine blades for rotation therewith. 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0007]     The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which:  
         [0008]      FIG. 1  is a schematic cross-sectional view of a gas turbine engine constructed in accordance with an aspect of the present invention;  
         [0009]      FIG. 2  is an enlarged view of a portion of the gas turbine engine of  FIG. 1 ; and  
         [0010]      FIG. 3  is a view taken along lines  3 - 3  of  FIG. 2 . 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0011]     Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,  FIG. 1  illustrates a representative gas turbine engine, generally designated  10 . The engine  10  has a longitudinal center line or axis A and an outer stationary annular casing  12  disposed concentrically about and coaxially along the axis A. The engine  10  has a high-pressure compressor  14 , combustor  16 , and high pressure turbine (“HPT”)  18  arranged in serial flow relationship, collectively forming a core  20 . A forward compressor rotor (e.g., low-pressure compressor, fan, or booster)  22  may be provided, driven by a low-pressure turbine  24  through an LP shaft  26 . In operation, pressurized air from the compressor  14  is mixed with fuel in the combustor  16  and ignited, thereby generating combustion gases. Some work is extracted from these gases by the high pressure turbine  18  which drives the compressor  14  via shaft  28 , and by the low pressure turbine  24 , which drives the booster  22 . The combustion gases then flow into an aft fan assembly  30  disposed aft of the core  20 . The aft fan assembly  30  comprises a free turbine or work turbine  32  which drives an integral aft fan  34 .  
         [0012]      FIGS. 2 and 3  illustrate the aft fan assembly  30  in more detail. The aft fan assembly  30  includes a hub or rotor  36  carrying a plurality of compound blades  38  in dovetail slots  40  and extending radially therefrom. Each of the compound blades  38  includes a turbine blade  42 , an arcuate platform segment  44 , and a plurality of fan blades  46 .  
         [0013]     In the illustrated example, each compound blade  38 , including the turbine blade  42 , its platform segment  44 , and the associated fan blades  46  is made as an integral component, for example by casting, forging, machining, or by fabrication (e.g. welding, brazing) from sub-components. The compound blades  38  could also be built-up as a mechanical assembly of individual components.  
         [0014]     Each of the turbine blades  42  is an airfoil having a leading edge  48 , a trailing edge  50 , a tip  52 , a root  54 , a convex suction side  56 , and a concave pressure side  58 . The turbine blades  42  are shaped to extract energy from the stream of pressurized gases exiting the core  20  to turn the rotor  36 . Depending upon the particular application, the turbine blades  42  may be provided with internal channels (not shown) connected to a source of cooling air to lower their temperature.  
         [0015]     Each platform segment  44  extends away from the associated turbine blade  42  in axial and circumferential directions. The platform segments  44  abut each other and collectively define an annular platform  60  interconnecting the tips  52  of the turbine blades  42 .  
         [0016]     The fan blades  46  are grouped into circumferential arrays referred to as “rows” or “stages”. A row  62  of first fan blades  46 A extends radially outward from the platform  60 . Each of the first fan blades  46 A is an airfoil having leading and trailing edges, a tip and a root, and opposed pressure and suction sides.  
         [0017]     A row  72  of second fan blades  46 B extends radially outward from the platform  60 , downstream of the first fan blades  46 A. Each of the second fan blades  46 B is an airfoil having leading and trailing edges, a tip and a root, and opposed pressure and suction sides.  
         [0018]     The number of fan blades  46  in each row  62  and  72  will vary depending on the specific application. The fan blades  46  have a reduced chord as compared to prior art fan-on-turbine designs. In order to preserve a selected solidity ratio of the rows  62  and  72 , a greater number of fan blades  46 A and  46 B are used in each of the rows  62  and  72 , as compared to a prior art fan-on-turbine design. This results in each turbine blade  42  carrying two or more first fan blades  46 A and two or more second fan blades  46 B. In the illustrated example, three first fan blades  46 A and three second fan blades  46 B extend from each platform segment  44 , for a total of six fan blades  46  per turbine blade  42 . Greater or lesser numbers of fan blades  46  may be used for each turbine blade  42  to suit a specific application.  
         [0019]     The fan blades  46  are surrounded by an annular casing  82  having inner and outer walls  84  and  86 . The inner surface of the outer wall  86  defines the outer boundary of a bypass duct  88  and the outer surface of the inner wall  84  defines the inner boundary of the bypass duct  88 , in cooperation with the platform  60 . A circumferential array of airfoil-shaped fan stator vanes  90  extends radially inward into the bypass duct  88  between the first and second fan rows  62  and  72 , and serves to redirect air flow exiting the first fan blades  46 A into the second fan blades  46 B at a desired angle.  
         [0020]     A circumferential array of radially-extending, airfoil-shaped inlet guide vanes (“IGVs”)  92  may be disposed in the bypass duct  88  forward of the fan blades  46 . The IGVs  92 , or portions thereof, are moveable so as to change their effective angle of attack relative to the air flow entering the bypass duct  88 . The IGVs  92  may be adjusted during engine operation to modulate air flow through the aft fan  34 . The IGVs may be operated using appropriate actuators  94  under the control of a FADEC, PMC, manual control, or other known type of engine control (not shown).  
         [0021]     A circumferential array of radially-extending, airfoil-shaped outlet guide vanes (“OGVs”)  96  is also disposed in the bypass duct  88 , aft of the fan blades  46 .  
         [0022]     The above-described aft fan assembly  30  is able to achieve greater work input than prior art fan-on-turbine designs without adding the complexity of additional turbine stages. For example, if a single fan stage that is capable of producing a pressure ratio of about 2.0 at a design operating condition, the two-stage design described above could enable a pressure ratio of about 3.5. To the extent that enough energy is available from the turbine  32 , more stages of fan blades  46  could be added.  
         [0023]     The foregoing has described a high pressure ratio aft fan for a gas turbine engine. While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the spirit and scope of the invention. Accordingly, the foregoing description of the preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of illustration only and not for the purpose of limitation, the invention being defined by the claims.