Patent Abstract:
A rotor dual-blade for a gas turbine engine that has a first blade component extending radially between a root and a tip and a second blade component, separate from the first component, extending radially between a root and a tip, wherein the second blade component is downstream, in series, of the first blade component and at least the first blade component is made of metal while the second blade component is a light weight composite material.

Full Description:
TECHNICAL FIELD 
     The described subject matter relates generally to gas turbine engines, and more particularly to rotor blades. 
     BACKGROUND OF THE ART 
     Conventional gas-turbine engine blades are made of metal, such as titanium. Fan blades, of this type, are capable of withstanding the temperatures to which they are exposed, erosion resistance and have a relatively good chance of surviving impact with foreign bodies, such as bird strikes, without seriously affecting engine performance. Metal blades, however, are relatively heavy and so increase the overall weight of the engine and reduce its performance. Efforts have been made, therefore, in recent years to develop blades made of alternative, lighter materials such as composite materials, for example, carbon fiber epoxy composites. The problem with such composite blades is that they are not as robust as metal blades and are more easily damaged by contact with foreign objects. Attempts have been made to protect the leading edge of the blades, which are most likely to be subject to damage, by means of metal sheaths. Examples of composite blades are described in, U.S. Pat. Nos. 5,881,972 and 7,896,619. 
     Accordingly, there is a need to provide an improved blade for gas turbine engines. 
     SUMMARY 
     In one aspect there is provided a rotor blade for a gas turbine engine comprising a first airfoil component extending radially between a root and a tip and a second airfoil component, separate from the first airfoil component, extending radially between a root and a tip, wherein the second airfoil component is downstream, in series, of the first airfoil component and at least the first airfoil component is made of metal. 
     In a another aspect there is provided a fan for a turbo fan engine comprising an array circumferentially spaced transonic dual rotor blades with each dual rotor blade having a leading blade component and a separate trailing blade component arranged in series with the leading blade component; the leading blade component made of a metal to resist to foreign object damage and erosion, and the trailing blade component made of a relatively lighter material providing enhanced aerodynamic characteristics to the dual rotor blade such that the weight of the dual rotor blade is less than a similar rotor blade made of solid metal. 
     In a further aspect there is provided a method of forming a rotor blade for a gas turbine engine comprising the steps of forming an annular base, mounting an array of dual rotor blade assemblies in a circumferential spaced apart arrangement on said base which each rotor blade assembly extending radially between a root on the base and a tip, the improvement including the steps of arranging in each dual rotor blade assembly a leading blade component selected from a metal and a trailing blade component selected from a suitable composite material suitable to reduce the weight of the rotor assembly, in series relative to the leading blade component and adjusting the trailing rotor component to tune the dual rotor blade. 
     In a still further aspect there is provided a method of forming a rotor blade for a fan in a turbo fan turbine engine comprising the steps of forming an annular hub, mounting an array of rotor blade assemblies in a circumferentially spaced apart arrangement on said hub which each rotor blade assembly extending radially between a root on the hub and a tip; the improvement including the steps of arranging in each rotor blade assembly a leading blade component selected from a metal suitable to withstand foreign object damage at transonic tip speeds and a trailing blade component selected from a suitable composite material suitable to reduce the weight of the rotor assembly, in series relative to the leading blade component and adjusting the trailing rotor component to tune the rotor blade. 
    
    
     
       DESCRIPTION OF THE DRAWINGS 
       Reference is now made to the accompanying figures in which: 
         FIG. 1  is a schematic cross-sectional view of a fan type gas turbine engine; 
         FIG. 2  is an axial cross-section of a portion of the engine with a side view of a detail of a preferred embodiment; 
         FIG. 3  is a schematic cross section of the detail shown in  FIG. 2 , at right angles thereto, showing a particular feature thereof; and 
         FIG. 4  is a schematic cross section of the detail shown in  FIG. 2 , at right angles thereto, showing a different feature compared to  FIG. 3 . 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically depicts a turbofan engine A which, as an example, illustrates the application of the described subject matter. The turbofan engine A includes a nacelle  10 , a low pressure spool assembly which includes at least a fan  12  and a low pressure turbine  14  connected by a low pressure shaft  16 , and a high pressure spool which includes a high pressure compressor  18  and a high pressure turbine  20  connected by a tie-shaft  22  and a high pressure shaft  24 . The engine further comprises a combustor  26 . 
     Advanced fans in turbofan engines are transonic with high rotor tip speeds. These transonic fans have to be strong enough to take a certain size of foreign object damage without a significant performance loss. At the same time, the reduction of weight is a key requirement for aircraft engine design. 
     The fan  12  described in the following example combines the features of transonic rotor tip speed with relatively lower weight compared to fans with existing all-titanium fan blades. The fan  12  includes an array of circumferentially spaced dual rotor blades, each made up of leading rotor blade  32  and trailing rotor blade  34 . The leading rotor blade  32  is made of a strong metal such as titanium or stainless steel. Other equivalent or superior materials may also be contemplated, as long as the criteria of resistance to foreign object damage and resistance to erosion are maintained. The term “metal” is defined herein to include such equivalent or superior materials. 
     The trailing rotor blade  34  is constructed of a lighter composite material. In one embodiment the composite material comprises carbon nanotubes. In another embodiment, carbon fibers are placed in multiple layers and are embedded with a polymer resin such as an epoxy-based resin. The trailing rotor blade  34  has the function of enhancing the aerodynamic characteristics of the dual fan blade while reducing the weight coefficient of the combined leading rotor blade  32  and the trailing rotor blade  34  (dual fan blade). 
     Each rotor blade  32 ,  34  includes a root  36   a ,  36   b  respectively. The roots  36   a ,  36   b  may be combined or separate. The leading rotor blade  32  has a tip  38  while the trailing rotor blade  34  has a tip  40 . The dual rotor blade  32 ,  34  includes a leading edge  42  and a trailing edge  44 . 
     A lengthwise gap  48  is defined between the leading rotor blade  32  and the trailing rotor blade  34 . The gap is quite small (exaggerated in the drawings) and will generally be in the range of 1% to 5% of the blade pitch. The gap may be filled with an elastomer such as rubber. 
     The fan  12 , in the present embodiment, has a weight advantage over a conventional metal fan, while at the same time having the Foreign Object Damage resistance of a metal fan because the leading rotor blade  32  covers the impacted region of Foreign Object Damage. From an aerodynamic standpoint, the dual rotor blades  32  and  34 , have the further advantage of producing lower pressure losses than a single rotor blade with the aerodynamic loading or turning. Further advantages of the fan  12  include that wakes produced thereby may be weaker than those of the equivalent single rotor design, and as such will reduce the fan noise. The fan noise may be further reduced by optimising the loading balance between the leading rotor blade  32  and the trailing rotor blade  34 . 
     Fan flutter is a challenging design issue for transonic fans. The dual fan concept that is described herein provides a further degree of freedom to tune the leading, upstream rotor blade  32  and the trailing, downstream rotor blade  34 . 
       FIG. 3  and  FIG. 4  illustrate alternate designs with the type of tuning that may be possible.  FIG. 3  illustrates the clocking of the trailing rotor blade  34  where arrow  50  illustrates the possible circumferential movement of the trailing rotor blade  34  relative to the leading rotor blade  32 , to change the relative angular position of the trailing rotor blade  34  with regard to the leading rotor blade  32 . 
     Similarly in  FIG. 4 , arrow  60  illustrates the axial movement that the trailing rotor blade  34  can achieve such that the trailing rotor blade  34  can overlap with the leading rotor blade  32  thus reducing the chord length of the dual blade while changing the angle thereof. 
     Thus the trailing rotor blade  34  may act as an aileron relative to the leading rotor blade  32  and may be tuned for ultimate aerodynamic performance. 
     The dual blade concept is shown as a fan, described herein above. However it is contemplated that the same concept may be applied to compressor or turbine rotors. 
     The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the described subject matter. Still other modifications which fall within the scope of the described subject matter will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.

Technology Classification (CPC): 5