Abstract:
A method for reducing vortex promotion of a rotor blade used in high core PR fans comprises shifting a curvature distribution of a blade section immediately adjacent to the hub such that a major turning of the suction side of the section is located near a trailing edge of the blade.

Description:
FIELD OF THE INVENTION 
     The present invention relates to an aircraft gas turbine engine, and more particularly to a rotor blade of an aircraft gas turbine engine, with an enhanced profile. 
     BACKGROUND OF THE INVENTION 
     A boost-less gas turbine engine does not include a boost compressor stage and therefore typically includes a high core pressure ratio (PR) fan which is adapted to compensate for the missing boost stage. 
     Conventional high core PR fan blades are usually configured with extreme blade turning immediately above the fan hub, which creates a very acute local angle between the blade suction side and the fan hub towards the blade trailing edge. Such an acute angle can help create and/or worsen a corner vortex at the trailing edge of the blade, potentially adversely affecting the quality of airflow at the hub area feeding into downstream blade rows of the compressor, and thereby reducing the overall engine efficiency and stability. Efforts have been made to solve this problem. For example U.S. Pat. No. 6,331,100 teaches providing an S-bowed stacking axis along which centers of gravity of the sections of the blade are aligned, in order to permit the trailing edge to be oriented substantially normal to the root of the bowed suction side and to lean hindward thereabove. U.S. Pat. No. 6,299,412 teaches that the airfoil suction side is laterally or tangentially bowed along the trailing edge near or adjacent the root at the intersection with the disk perimeter in order to increase blade efficiency and improve stall margin. 
     Nevertheless, there is still a need for improved approaches and solutions to better solve the corner vortex problem. 
     SUMMARY OF THE INVENTION 
     One object of the present invention is to provide a fan blade of an aircraft gas turbine engine, with an enhanced profile. 
     In accordance with one aspect of the present invention, there is a method provided for reducing vortex promotion of a rotor blade used in a high core pressure ratio fans. The blade has a plurality of sections extending from a hub to a tip thereof. The method comprises providing a curvature distribution to a first blade section immediately adjacent to the hub such that a major turning of a suction side of the respective sections is located near a trailing edge of the blade, thereby increasing an angle of the section at a suction side between the trailing edge and a periphery of the hub. 
     A predetermined total curvature turning of the suction side of each section of the rotor blade is preferably predetermined and is unchanged in the step of providing the curvature distribution. 
     In accordance with another aspect of the present invention, there is a rotor blade of a gas turbine engine affixed to a hub, which comprises a plurality of sections extending from the hub to a tip of the blade, defining leading and trailing edges extending between the hub and tip thereof, and pressure and suction sides joining at the respective leading and trailing edges. The respective sections define different curvature distributions at the suction side thereof to create a major turning of the suction side of a first blade section immediately adjacent to the hub. The major turning of the suction side is located near the trailing edge relative to the remaining sections. 
     In accordance with a further aspect of the present invention, there is a high core pressure ratio fan of a gas turbine engine which comprises a rotor having a hub, and a plurality of blades extending from the hub. Each blade defines leading and trailing edges extending from the hub to a tip thereof, and pressure and suction sides extending between the hub and the tip and adjoining at the respective leading and trailing edges. Each blade includes a plurality of sections thereof with different curvature distributions at the suction side, thereby forming a curved surface of the suction side of the blade having a pocket located in an area of the suction side immediately adjacent to the hub in a vicinity of the trailing edge relative to the remaining area of the suction side, thereby causing the trailing edge to have a turning portion immediately adjacent to the hub adapted to increase an angle of the trailing edge at the suction side relative to a periphery of the hub. 
     The present invention advantageously achieves the required high core PR profile of a high core pressure ratio fan while minimizing the promotion of a corner vortex, thereby improving the airfoil flow quality of the fan, and thus improving an overall efficiency and stability of the engine performance. 
     Other features and advantages of the present invention will be better understood with reference to a preferred embodiment described hereinafter. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Reference will now be made to the accompanying drawings, showing by way of illustration the preferred embodiment thereof, in which: 
         FIG. 1  is a schematic cross-sectional view of a turbofan gas turbine engine for use in aircraft, illustrating one application of the present invention; 
         FIG. 2  is partial perspective view of the high core PR fan used in the turbofan gas turbine engine, incorporating one embodiment of the present invention; 
         FIG. 3  is a partial edge-on view of the trailing edge of a conventional fan blade with the remainder thereof being omitted in the interests of clarity, showing the trailing edge of a conventional fan blade having a nominal curvature distribution in a radial plane; 
         FIG. 4  illustrates in respective axial planes, the profiles of a number of sections of the conventional fan blade of  FIG. 3 ; 
         FIG. 5  is a partial edge-on view of the trailing edge of a fan blade having a shifted curvature distribution in a radial plane according to the embodiment of  FIG. 2 ; and 
         FIG. 6  illustrates in respective axial planes, profiles of the first four sections of the fan blade of  FIG. 5 . 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
     A turbofan gas turbine engine illustrated schematically in  FIG. 1  incorporates an embodiment of the present invention, presented as an example of the application of the present invention, and includes a housing or nacelle  10 , a low pressure spool assembly seen generally at  12  which includes a high core pressure ratio fan  14 , low pressure compressor  16  and low pressure turbine  18 , and a high pressure spool assembly seen generally at  20  which includes a high pressure compressor  22  and a high pressure turbine  24 . There is provided a burner seen generally at  25  which includes an annular combustor  26  and a plurality of fuel injectors  28  for mixing liquid fuel with air and injecting the mixed fuel/air flow into the annular combustor to be ignited for generating combustion gases. The high core pressure ratio fan  14  can also be used in other types of turbofan gas turbine engines, for example, a boost-less gas turbine engine which does not include the low pressure compressor  16 . 
     In  FIG. 2 , a portion of the high core pressure ratio fan  14  is illustrated and includes an annular hub  32  and a plurality of circumferentially spaced apart rotor blades or airfoils  34  extending radially outwardly from a periphery  31  of the annular hub  32 . The rotor blade  34  includes a leading edge  38  and a trailing edge  40 , with respect to the airflow direction as indicated by arrow  42  passing through and being compressed by the rotor blades  34 . The leading and trailing edges  38 ,  40  extend from the periphery  31  of the hub  32  to a tip  36  of the rotor blade  34 . The rotor blade  34  further includes a concave pressure side  44  and a convex suction side  46 , joining together at the respective leading and trailing edges  38 ,  40 . The rotor blade  34  rotates in a rotational direction as illustrated by arrow  48 . The pressure side  44  and the suction side  46  are aerodynamically configured for maximizing the efficiency of air compression and achieve desired pressure ratio. For design purposes, the rotor blade  34  is divided into a number of sections S 1 , S 2 , S 3 , S 4  to Sn stacked finite radially and outwardly from the hub  32  to the tip  36  of the rotor blade  34 . It will all be understood that the sections, as illustrated in  FIGS. 2 and 5  are enlarged for the purposes of description only. 
     Referring to FIGS.  2  and  5 – 6 , the rotor blade  34  of the high core pressure ratio fan  14  includes a high blade turning in a radial plane, in the sections immediately above the fan hub  32 , as illustrated by sections S 1 , S 2  and S 3  in  FIG. 2 , and also defines a trailing edge angle A at the suction side  46  with respect to the periphery  31  of the hub  32 , which would have been much more acute than is illustrated in  FIGS. 2 and 5  if the blade was conventionally designed. This will be further discussed with reference to  FIGS. 3 and 4 . 
       FIG. 3  illustrates a prior art high core pressure ratios rotor blade  34   a  and an annular hub  32   a  of a conventional high core pressure ratio fan  14   a , and having a trailing edge  40   a  with angle A 1 . The profiles of sections S 1   a , S 2   a , S 3   a  and S 4   a  are illustrated in  FIG. 4  in which B 1  indicates an angle of the suction side of the rotor blade  34   a  at the leading edge with respect to a plane parallel to the longitudinal axis. Similarly, B 2  indicates an angle of the suction side of the rotor blade  34   a  at the trailing edge  40   a  with respect to a plane parallel to the longitudinal axis. Different sections S 1 –S 4 , etc. of the conventional rotor blade  34   a  have different angles B 1  and B 2 . A total blade turning of the suction side curvature of each section of rotor blade  34   a  is determined by angles B 1  and B 2 . 
     Again referring to  FIGS. 2 ,  5  and  6 , in accordance with the present invention the curvature distribution in an axial plane, of the individual section of the rotor blade  34  immediately adjacent to the annular hub  32  (for example section S 1 ), is provided with a major turning portion  50  of the suction side curvature near the trailing edge  40 . This section is referred to as a “back-loaded section”. Sections adjacent this “back loaded” section (for example, S 2 ) are preferably “front loaded”, relative to S 1  such that the turning  52  of S 2  is nearer the leading edge. This front-loading (relative to S 1 ) can be applied to  52 ,  53 , etc. as required, along with changes for the back-loading to S 1  to “open up” angle A, i.e. to make angle A as large as possible until it is sufficient to alleviate the corner vortex formation for the given fan blade being designed. 
     The fan blade  34  is preferably configured with sections each defining a major turning portion of the suction side curvature thereof in an axial plane. These major turning portions defined by the respective sections of the rotor blade  34  are preferably positioned in a sequence gradually approaching the leading edge  38  and the tip  36  of the rotor blade  34 . Thus, the suction side  46  of the rotor blade  34  forms a pocket  54  located in an area immediately adjacent to the hub  32  in a vicinity of the trailing edge  40  relative to the remaining area of the pressure side  44 , thereby causing the trailing edge  40  to have a turning portion immediately adjacent to the hub  32  in order to increase the angle A of the trailing edge  40  at the suction side  46  relative to the periphery  31  of the hub  32 . 
     In a comparison of  FIG. 5  with  FIG. 3 , it is apparent that the trailing edge  40  of the rotor blade  34  represents a curved line having a tendency to increase the angle A (relative to A 1  of  FIG. 3 ) of the section S 1  at the suction side  46  between the trailing edge  40  and the periphery  31  of the hub  32 . A portion of the trailing edge  40  of the rotor blade  34  immediately adjacent to the hub  32  and extending therefrom preferably tends toward a normal intersection with the hub  32 . In contrast, however, the trailing edge  40   a  of the prior art rotor blade  34   a  shown in  FIG. 3 , has a portion thereof immediately adjacent to the hub  32   a  and extending therefrom at a relatively acute angle, as indicated by A 1 . Although a substantially normal intersection is shown in  FIG. 5 , according to the present invention any increase in angle A, such that A&gt;A 1 , offers improvement over the prior art. 
     It should be noted that shifting the curvature distribution of the suction side of individual sections of the rotor blade  34  should not change a total curvature turning of the suction side  46  of the rotor blade  34 . The total curvature turning is nominal or predetermined in order to achieve a required high core pressure ratio of the fan  14 . For example, the rotor blade  34  is designed to achieve the substantially same high core pressure ratio which the conventional rotor blade  34   a  of  FIGS. 3–4  is configured to achieve. Thus, the total curvature turning of the suction side of the individual sections such as S 1 , S 2 , etc. should be substantially equal to the total curvature turning of the respective sections of conventional rotor blade  34   a , such as S 1   a , S 2   a , etc. This condition can be assured when the angle B 1  and B 2  of the individual sections (only shown with section S 1 ) at leading and trailing edges  38 ,  40  of the rotor blade  34 , are substantially equal to the angles B 1 , B 2  of the corresponding individual sections (only shown with section S 1   a ) at the leading and trailing edges of the prior art rotor blade  34   a  of  FIG. 4 . 
     It should also be noted that the drawings are schematical and are exaggerated to more clearly illustrate the present invention but are not intended to illustrate a proportional physical structure of the embodiment of the present invention. 
     Modifications and improvements to the above-described embodiment of the present invention may become apparent to those skilled in the art. For example, the present invention may be employed with removably bladed fan rotors or integrally-bladed rotors, and with blades of any profile or sweep angle. The foregoing description is intended to be exemplary rather than limiting. The scope of the present invention is therefore intended to be limited solely by the scope of the appended claims.