Abstract:
A leading edge structure for use in an aerospace vehicle includes a body having a flowpath surface which defines a leading edge adapted to face an air flow during operation, and an opposed inner surface. The body is segmented into a plurality of portions having varying thermal properties and/or mechanical discontinuities, so as to promote stress concentrations in ice attached to the flowpath surface.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    This invention relates generally to aerospace vehicle structures and more particularly to designs for improving ice shedding characteristics from such structures. 
         [0002]    All aircraft include various “leading edge structures”, i.e. exposed surfaces that face the direction of flight. These surfaces include, for example, parts of the fuselage, wings, control surfaces, and powerplants. 
         [0003]    One common type of aircraft powerplant is a turbofan engine, which includes a turbomachinery core having a high pressure compressor, combustor, and high pressure turbine in serial flow relationship. The core is operable in a known manner to generate a flow of propulsive gas. A low pressure turbine driven by the core exhaust gases drives a fan through a shaft to generate a propulsive bypass flow. The low pressure turbine also drives a low pressure compressor or “booster” which supercharges the inlet flow to the high pressure compressor. 
         [0004]    Certain flight conditions allow for ice build up on the leading edge structures, and in particular the fan and booster flowpath areas of the engine. These areas include the blades, spinner cone, and static vane and fairing leading edges. The FAA requires certification testing at these flight points to demonstrate the ability to maintain engine thrust once the ice sheds from the various components and ingests into the engine. 
         [0005]    One particular leading edge structure of interest is the engine&#39;s fan splitter. The splitter is an annular ring with an airfoil leading edge that is positioned immediately aft of the fan blades. Its function is to separate the airflow for combustion (via the booster) from the bypass airflow. It is desired for the splitter and other leading edge structures to have mechanical, chemical, and thermal properties such that ice build up and shed volume is minimized during an icing event. This in turn minimizes risk of compressor stall and compressor mechanical damage from the ingested ice. 
         [0006]    Prior art turbofan engines have splitters made from titanium, which is known to provide favorable ice shed properties. The downside of titanium is the expense and weight when compared to conventionally treated aluminum. However, conventionally treated aluminum is believed to behave poorly in an aircraft icing environment. Examples of conventionally treated aluminum include but are not limited to chemical conversion coatings and anodization. 
         [0007]    Leading edge structures can also be protected with known coatings that are referred to as “icephobic” or “anti-ice” coatings, for example polyurethane paint or other organic coatings. These coatings have the effect of lowering adhesion forces between ice accretions and the protected component. While these coatings can improve ice shedding characteristics, their erosion resistance may be not adequate to protect leading edge structures from the scrubbing effect of airflows with entrained abrasive particles which are encountered in flight. 
       BRIEF SUMMARY OF THE INVENTION 
       [0008]    These and other shortcomings of the prior art are addressed by the present invention, which provides components having icephobic plating that reduces and/or modifies ice adhesion forces to promote ice release and reduce shedding of large ice pieces. 
         [0009]    According to one aspect, the invention provides a leading edge structure for use in an aerospace vehicle, including: (a) a body having a flowpath surface which defines a leading edge adapted to face an air flow during operation; and (b) a plurality of mechanical discontinuities formed in the flowpath surface, the mechanical discontinuities adapted to promote stress concentrations in ice attached to the flowpath surface. 
         [0010]    According to another aspect of the invention, a splitter for a turbofan engine includes: (a) an annular body having a flowpath surface which defines a leading edge adapted to face an air flow during operation; and (b) a plurality of mechanical discontinuities formed in the flowpath surface, the mechanical discontinuities adapted to promote stress concentrations in ice attached to the flowpath surface. 
         [0011]    According to another aspect of the invention, a leading edge structure for use in an aerospace vehicle includes a body having a flowpath surface which defines a leading edge adapted to face an air flow during operation, and an opposed inner surface. The body is segmented into a plurality of portions having varying thermal properties, so as to promote stress concentrations in ice attached to the flowpath surface. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS  
         [0012]    The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which: 
           [0013]      FIG. 1  is a perspective view of an aircraft powered by a pair of high-bypass turbofan engines, incorporating icing-resistant components constructed according to an aspect of the present invention; 
           [0014]      FIG. 2  is a schematic half-sectional view of an engine shown in  FIG. 1 ; 
           [0015]      FIG. 3  is a half -sectional view of a splitter shown in  FIG. 2 ; 
           [0016]      FIG. 4  is a view taken from forward looking aft at the splitter of  FIG. 3 ; 
           [0017]      FIG. 5  is a view taken along lines  5 - 5  of  FIG. 4 ; 
           [0018]      FIG. 6A  is a view taken along lines  6 - 6  of  FIG. 5 ; 
           [0019]      FIG. 6B  is a forward looking aft view of a variation of the splitter of  FIG. 6A ; 
           [0020]      FIG. 7  is a view taken from forward looking aft at an alternative splitter; 
           [0021]      FIG. 8  is a taken along lines  8 - 8  of  FIG. 7 ; 
           [0022]      FIG. 9A  is a view taken along lines  9 - 9  of  FIG. 8 ; 
           [0023]      FIG. 9B  is a forward looking aft view of a variation of the splitter of  FIG. 9A ; 
           [0024]      FIG. 10  is a view taken from forward looking aft at another alternative splitter; 
           [0025]      FIG. 11  is a taken along lines  11 - 11  of  FIG. 10 ; 
           [0026]      FIG. 12A  is a view taken along lines  12 - 12  of  FIG. 11 ; and 
           [0027]      FIG. 12B  is a forward looking aft view of a variation of the splitter of  FIG. 12A   
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0028]    Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,  FIG. 1  depicts a known type of commercial aircraft  10  which includes a generally tubular fuselage  12 , wings  14  carrying turbofan engines  16  mounted in nacelles  18 , and an empennage comprising horizontal and vertical stabilizers  20  and  22 . Each of these components includes one or more exposed surfaces having a curved or airfoil-like cross-section that faces the direction of flight (in other words an aerodynamic leading edge). These surfaces are referred to herein as “leading edge structures”. While the present invention will be described further in the context of a gas turbine engine, it will be understood that the principles contained therein may be applied to any type of leading edge structure. 
         [0029]    As shown in  FIG. 2 , the engine  16  has a longitudinal axis “A” and includes conventional components including a fan  24 , a low pressure compressor or “booster”  26  and a low pressure turbine (“LPT”)  28 , collectively referred to as a “low pressure system”, and a high pressure compressor (“HPC”)  30 , a combustor  32 , and a high pressure turbine (“HPT”)  34 , collectively referred to as a “gas generator” or “core”. Various components of the nacelle  18  and stationary structures of the engine  16 , including a core nacelle  36 , cooperate to define a core flowpath marked with an arrow “F”, and a bypass duct marked with an arrow “B”. 
         [0030]    A stationary annular splitter  38  (also seen in  FIG. 3 ) is positioned at the forward end of the core nacelle  36 , between the bypass duct B and the core flowpath F. The flowpath surface  40  of the splitter  38  includes a radially-outward-facing portion  41  and a radially-inward-facing portion  43 . The two portions are demarcated by an aerodynamic leading edge  39 . An inner surface  45 , not exposed to the primary flowpath, is disposed opposite the flowpath surface  40 . The splitter  38  is an example of a leading edge structure as described above. The splitter  38  may be a single continuous ring, or it may be built up from arcuate segments. 
         [0031]    The flowpath surface  40  includes one or more discontinuities for the purpose of improving ice shed characteristics. As shown in  FIGS. 3-6A , the splitter  38  has a radial array of generally axially aligned grooves  42  formed therein. As an example, the width “W” of the grooves may be from as small as about 0.38 mm (0.015 in.) up to as large as 50% of the circumference of the splitter  38 .  FIG. 6B  illustrates a slightly different splitter  38 ′ in which the flowpath surface  40 ′ has grooves  42 ′ that are curved. They may be curved so as to be parallel to the local flowfield during operation.  FIGS. 7-9A  illustrate an alternative splitter  138  that has a radial array of generally axially aligned, raised ribs  142  protruding from its flowpath surface  140 . The spacing “S” of the grooves  42  or ribs  142  in the circumferential direction may be selected to cause ice to breakup into relatively small pieces. As an example, about 24 to about 140 features distributed around the circumference are believed to be suitable for this purpose.  FIG. 9B  illustrates a slightly different splitter  138 ′ in which the flowpath surface  140 ′ has ribs  142 ′ that are curved. They may be curved so as to be parallel to the local flowfield during operation. Various patterns of grooves or ribs running in different directions (axial, circumferential, and combinations of each direction etc.) may be used. 
         [0032]      FIGS. 10-12A  illustrate another alternative splitter  238  whose flowpath surface  240  includes alternating sections  242 A and  242 B having substantially different thicknesses such that adjacent sections are offset in a direction normal to the flowpath surface (i.e. in the radial direction in illustrated example). The delineations between adjacent sections  242 A and  242 B present generally radially aligned faces  244  which act as discontinuities in the flowpath surface  240 .  FIG. 12B  illustrates a slightly different splitter  238 ′ in which the flowpath surface comprises segments  242 ′ that are tapered in thickness in the circumferential direction. The delineations between adjacent sections  242 ′ present generally radially curved faces  244 ′ which act as discontinuities. The faces  244 ′ may be curved so as to be parallel to the local flowfield during operation. As with the grooves or ribs, the delineations may be implemented in various patterns running in different directions (axial, circumferential, etc.) 
         [0033]    In operation, the engine  10  will be exposed to icing conditions, namely the presence of moisture in temperatures near the freezing point of water. Ice will naturally tend to form on the leading edge structures including the splitter  38 . As the ice mass builds up, it protrudes into the air flow and increasing aerodynamic (drag) forces act on it, eventually causing portions of it to shed from the splitter  38 . The presence of the discontinuities described above promotes stress concentrations and introduces mechanical stresses into the ice. The result is that pieces of the ice break off and shed downstream when they are a smaller size than would otherwise be the case. This avoids excessive cooling and foreign object damage in the high pressure compressor  30 . 
         [0034]    In addition to, or as an alternative to the techniques described above, the thermal properties of the leading edge structure can be varied by changes in either alloy type or thickness. Changes to surface properties and texture may also help with heat transfer. Also, the internal (non-flowpath) surfaces can be varied in order to achieve the desired thermal variations. For example, the local thickness variation described above can be achieved by adding thickness to the inner surface, while leaving the flowpath surface unchanged). 
         [0035]    The foregoing has described aerospace structures adapted for improved ice shedding characteristics. 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.