Abstract:
A method for assembling a turbine engine to facilitate preventing ice accumulation on the turbine engine during engine operation, the gas turbine engine including a fan assembly. The method includes coupling a plurality of fan blades to a fan rotor, and applying a polyurethane material to at least a portion of at least one of the fan blades to facilitate preventing ice accumulation on the at least one fan blade.

Description:
BACKGROUND OF THE INVENTION  
       [0001]     This invention relates generally to gas turbine engines, and more particularly, to methods and apparatus for operating gas turbine engines.  
         [0002]     Gas turbine engines typically include an inlet, a fan, low and high pressure compressors, a combustor, and at least one turbine. The compressors compress air which is channeled to the combustor where it is mixed with fuel. The mixture is then ignited for generating hot combustion gases. The combustion gases are channeled to the turbine(s) which extracts energy from the combustion gases for powering the compressor(s), as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator.  
         [0003]     When engines operate in icing conditions, i.e., exposed to clouds of super-cooled water droplets, ice may accumulate on exposed engine structures. More specifically, if engines are operated within icing conditions at low power for extended periods of time, ice accumulation within the engine and over exposed engine structures may be significant. Over this time, large quantities of ice can accrete and may cause high engine vibrations until the ice is shed. This ice may enter the high pressure compressor. Such a condition, known as a shed cycle, may cause the compressor discharge temperature to be suddenly reduced. In response to the sudden decrease in compressor discharge temperature, the corrected core speed increases in the aft stages of the high pressure compressor. This sudden increase in aft stage corrected core speed may adversely impact compressor stall margin. In some cases, it may also lead to compressor stall.  
         [0004]     To facilitate preventing ice accretion within the engine and over exposed surfaces adjacent the engine, at least some known engines include a control system that enables the engine to operate with an increased operating temperature and may include sub-systems that direct high temperature bleed air from the engine compressor to provide heat to the exposed surfaces. However, the increased operating temperature and the bleed systems may decrease engine performance. Such systems may also require valves to turn off the flow of the high temperature air during take-off and other high power operations to protect the engine. In addition to the increased cost, such valving may pose a reliability problem. As such, to further facilitate preventing ice accumulation, at least some known engines are sprayed with a deicing solution prior to operation. However, during flight, de-icing solutions are not applied. More specifically, during flight, evaporative cooling may still cause freezing and ice accumulation over external engine surfaces, such as a front frame of the engine. Conventional electrical heating is an option, but it requires electricity for performing the de-icing operation and may require relatively large generators at low speed conditions, electrical circuits, and complex interactive logic with the airplane&#39;s computers with the attendant increased cost, weight and performance penalties.  
       BRIEF SUMMARY OF THE INVENTION  
       [0005]     In one aspect, a method for assembling a turbine engine is provided. The method includes coupling a plurality of fan blades to a fan rotor, and applying a polyurethane material to at least a portion of at least one of the fan blades to facilitate preventing ice accumulation on the at least one fan blade.  
         [0006]     In another aspect, a fan assembly for a gas turbine engine is provided. The fan assembly includes a fan rotor disk, a plurality of fan blades coupled to the fan rotor disk, and a polyurethane material applied to at least a portion of at least one of the fan blades to facilitate preventing ice accumulation on the at least one the fan blade.  
         [0007]     In a further aspect, a gas turbine engine assembly is provided. The gas turbine engine assembly includes a core gas turbine engine, and a fan assembly coupled to the core gas turbine engine. The fan assembly includes a fan rotor disk, a plurality of fan blades coupled to the fan rotor disk, the plurality of fan blades fabricated at least partially from a titanium material, and a polyurethane material applied to at least a portion of at least one of the fan blades to facilitate preventing ice accumulation on the at least one the fan blade. 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0008]      FIG. 1  is a schematic illustration of an exemplary gas turbine engine;  
         [0009]      FIG. 2  is a perspective view of the fan assembly shown in  FIG. 1 :  
         [0010]      FIG. 3  is a forward view of the exemplary fan blade shown in  FIG. 2 ;  
         [0011]      FIG. 4  is a side elevation view of the fan blade shown in  FIG. 3 ; and  
         [0012]      FIG. 5  is a sectional view of the fan blade shown in  FIG. 3 . 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0013]      FIG. 1  is a schematic illustration of an exemplary gas turbine engine assembly  10  having a longitudinal axis  11 . Gas turbine engine assembly  10  includes a fan assembly  12 , a high pressure compressor  14 , and a combustor  16 . Engine  10  also includes a high pressure turbine  18 , a low pressure turbine  20 , and a booster  22 . Fan assembly  12  includes an array of fan blades  24  extending radially outward from a rotor disk  26 . Engine  10  has an intake side  28  and an exhaust side  30 . In the exemplary embodiment, the gas turbine engine is a GE90 available from General Electric Company, Cincinnati, Ohio. Fan assembly  12 , booster  22 , and turbine  20  are coupled together by a first rotor shaft  31 , and compressor  14  and turbine  18  are coupled together by a second rotor shaft  32 .  
         [0014]     In operation, air flows through fan assembly  12  and compressed air is supplied to high pressure compressor  14  through booster  22 . The booster discharge air is further compressed and delivered to combustor  16 . Hot products of combustion (not shown in  FIG. 1 ) from combustor  16  are utilized to drive turbines  18  and  20 , and turbine  20  is utilized to drive fan assembly  12  and booster  22  by way of shaft  31 . Engine  10  is operable at a range of operating conditions between design operating conditions and off-design operating conditions.  
         [0015]     In the exemplary embodiment, a bypass duct  40  is utilized to bypass a portion of the airflow from fan assembly  12  around gas turbine engine  10 . More specifically, bypass duct  40  is defined between an outer casing  42  and a splitter  44  that substantially circumscribes booster  22 . Accordingly, a first portion of the airflow compressed by fan assembly  12  is divided between bypass duct  40  and an inlet  46  to the booster  22  utilizing splitter  44  coupled upstream from gas turbine engine  10 .  
         [0016]      FIG. 2  is a perspective view of fan assembly  12  shown in  FIG. 1 .  FIG. 3  is an aft looking forward view of fan blade  24  shown in  FIG. 2 .  FIG. 4  is a side elevation view of fan blade  24 .  FIG. 5  is a sectional view of fan blade  24 . Each fan blade  24  includes an airfoil  52  and an integral dovetail  54  that is used for mounting airfoil  52  to a rotor disk, such as rotor disk  26  (shown in  FIG. 1 ) in a known manner.  
         [0017]     Each airfoil  52  includes a first contoured sidewall  60  and a second contoured sidewall  62 . First sidewall  60  is convex and defines a suction side of airfoil  52 , and second sidewall  62  is concave and defines a pressure side of airfoil  52 . Sidewalls  60  and  62  are joined at a leading edge  64  and at an axially-spaced trailing edge  66  of airfoil  52 . More specifically, airfoil trailing edge  66  is spaced chordwise and downstream from airfoil leading edge  64 . First and second sidewalls  60  and  62 , respectively, extend longitudinally or radially outward in span from a blade root  68  positioned adjacent dovetail  54 , to an airfoil tip  70 . A dovetail platform  72  is positioned at blade root  68  and extends radially outward from first and second sidewalls  60  and  62 , respectively.  
         [0018]     The general configuration of each fan blade  24  may take any conventional form with or without the platform  72  or the dovetail  54 . For example, fan blade  24  may be alternatively formed integrally with the disk  26  as one assembly conventionally referred to as a blisk without a discrete and removable dovetail  54 . Fan blade  24  may also be of the conventional solid-type or hollow-type as desired. In the exemplary embodiment, each fan blade  24  is fabricated utilizing a metallic material such as, but not limited to, titanium. In an alternative embodiment, each fan blade  24  is fabricated utilizing a composite material.  
         [0019]     Moreover, although the invention is described herein with respect to fan blade  24 , it should be realized that the invention can be applied to any blades utilized within a gas turbine engine such as, but not limited to booster compressor blades.  
         [0020]     Accordingly, and in the exemplary embodiment, each fan blade  24  includes a material  100  that is affixed to at least a portion of fan blade  24  to facilitate shedding any ice that may form and/or accumulate on fan blade  24 . More specifically, and in the exemplary embodiment, material  100  is a material which has a lower ice adhesion characteristic than the parent material. For example, in the exemplary embodiment, a polyurethane material is affixed to second sidewall  62 , i.e. the pressure side of airfoil  52 . Polyurethane as used herein is defined as any polymer that includes a chain of organic units joined by urethane links.  
         [0021]     In the exemplary embodiment, material  100  has a width  110  that is between approximately 85% and approximately 95% of a width  112  of airfoil  52 . More specifically, material  100  is applied to airfoil  52  such that material  100  extends substantially from airfoil trailing edge  66  at least partially towards leading edge  64 . In the exemplary embodiment, material  100  is not carried to leading edge  64  due to the erosion seen at leading edge  64  and between approximately 5% and approximately 15% aft of leading edge  64 . Accordingly, and in the exemplary embodiment, material  100  has a width  110  that is between approximately 85% and approximately 95% of airfoil width  112 .  
         [0022]     Moreover, material  100  also has a length  120  that extends radially outwardly in span from blade root  68  at least partially towards airfoil tip  70 . In the exemplary embodiment, length  120  is selected based on the centrifugal load, i.e. the centrifugal force experienced by the ice accumulated on fan blade  24  during engine icing conditions. The centrifugal force is defined as a function of the mass of the ice, the radius at which the ice accumulates on fan blade  24 , and the tangential velocity of fan blade  24  at the specified radius. Accordingly, length  120  is pre-selected for each gas turbine engine based on the estimated force (I) required to dislodge any accumulated ice build-up on fan blade  24 . In the exemplary embodiment, length  120  is approximately one-third of a length  122  of airfoil  52 . In another exemplary embodiment, length  120  is approximately one-half of a length  122  of airfoil  52  for a fan operating at a slower speed.  
         [0023]     Thus, a first gas turbine engine operating within a first range of rotational speeds will include polyurethane material  100  having a length  120  affixed to each fan blade  24 . Whereas, a second gas turbine engine operating within a second range of rotational speeds, that is less than the first range of rotational speeds, will include polyurethane material  100  having a second length  126 , that is greater than length  120 , affixed to each fan blade  24  since the force available to dislodge the accumulated ice from fan blade  24  is reduced when the engine is operating at a reduced speed. Additionally, a third gas turbine engine operating within a third range of rotational speeds, that is greater than the first range of rotational speeds, will include polyurethane material  100  having a third length  124 , that is less than length  120 , affixed to each fan blade  24  since the force available to dislodge the accumulated ice from fan blade  24  is greater when the engine is operating at increased rotational speeds.  
         [0024]     During assembly, material  100  is affixed or applied to each fan blade  24 . More specifically, and in one embodiment, material  100  is a sprayed-on material  150  that applied to fan blade  24  using a typical painting process. In one embodiment, material  150  is applied to fan blade  24  such that material  150  is between approximately 3 mils, i.e. three one-thousandths of an inch, and approximately 10 mils thick. In the exemplary embodiment, material  150  has a thickness  160  that is approximately five mils. In another embodiment, material  100  is a sheet  170  of polyurethane material that is applied to fan blade  24 . In the exemplary embodiment, sheet  170  has a thickness  160  that is between approximately five mils and approximately twenty mils. During assembly, material  100  is tapered and/or feathered to form a relatively smooth transition from the metallic surface of blade  24  to the full thickness  160  of material  100  aft of leading edge  64  and at the full thickness  160  of material  100 .  
         [0025]     The above-described ice protection material is affixed to at least one gas turbine blade to facilitate shedding any ice that may accumulated on the gas turbine blade. In the exemplary embodiment, the ice protection material is applied to a plurality of gas turbine engine fan blades. The above-described ice-protection material is cost-effective and highly reliable in facilitating the prevention of ice accumulation along exposed surfaces of the engine. More specifically, the polyurethane ice protection material is applied to the fan blade pressure side as either a film or a spray on material to a width that is selected based on the G-load that the specific gas turbine fan blade is expected to realize. Accordingly, in the exemplary embodiment described herein, the polyurethane material is applied to the lower one-third of the pressure side of the blade and extends between approximately 85% and approximately 95% from the trailing edge of the fan blade towards the leading edge of the fan blade.  
         [0026]     As a result, an ice protection material is provided which facilitates reducing the adhesive strength of the ice which may form on the fan blade by approximately 50%. Thus ice forming on the fan blades will shed sooner with less residual unbalance between sheds compared to non-treated systems.  
         [0027]     Exemplary embodiments of an ice protection material are described above in detail. The ice protection material is not limited to the specific embodiments described herein, but rather, the polyurethane ice protection material may be applied to any portion of the gas turbine engine to facilitate ice shedding. For example, the ice protection material may be applied to portions of the booster compressor, splitter, and/or portions of the fan shroud.  
         [0028]     While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.