Abstract:
An unducted thrust producing system has a rotating element with an axis of rotation and a stationary element. The rotating element includes a plurality of blades, and the stationary element has a plurality of vanes configured to impart a change in tangential velocity of the working fluid opposite to that imparted by the rotating element acted upon by the rotating element. The system includes an inlet forward of the rotating element and the stationary element.

Description:
[0001]    This application is a national stage application under 35 U.S.C. §371(c) of prior-filed, co-pending, PCT application serial number PCT/US2013/066392, filed on Oct. 23, 2013, which claims priority to Provisional Patent Application Ser. No. 61/717,445 filed Oct. 23, 2012 and titled “PROPULSION SYSTEM ARCHITECTURE”, and is related to PCT application serial number PCT/US2013/066383, titled “UNDUCTED THRUST PRODUCING SYSTEM” filed on Oct. 23, 2013, and PCT application serial number PCT/US2013/066403, titled “VANE ASSEMBLY FOR AN UNDUCTED THRUST PRODUCING SYSTEM” filed on Oct. 23, 2013. All of the above listed applications are herein incorporated by reference. 
     
    
     BACKGROUND OF THE INVENTION 
       [0002]    The technology described herein relates to an unducted thrust producing system, particularly architectures for such systems. The technology is of particular benefit when applied to “open rotor” gas turbine engines. 
         [0003]    Gas turbine engines employing an open rotor design architecture are known. A turbofan engine operates on the principle that a central gas turbine core drives a bypass fan, the fan being located at a radial location between a nacelle of the engine and the engine core. An open rotor engine instead operates on the principle of having the bypass fan located outside of the engine nacelle. This permits the use of larger fan blades able to act upon a larger volume of air than for a turbofan engine, and thereby improves propulsive efficiency over conventional engine designs. 
         [0004]    Optimum performance has been found with an open rotor design having a fan provided by two contra-rotating rotor assemblies, each rotor assembly carrying an array of airfoil blades located outside the engine nacelle. As used herein, “contra-rotational relationship” means that the blades of the first and second rotor assemblies are arranged to rotate in opposing directions to each other. Typically the blades of the first and second rotor assemblies are arranged to rotate about a common axis in opposing directions, and are axially spaced apart along that axis. For example, the respective blades of the first rotor assembly and second rotor assembly may be co-axially mounted and spaced apart, with the blades of the first rotor assembly configured to rotate clockwise about the axis and the blades of the second rotor assembly configured to rotate counter-clockwise about the axis (or vice versa). In appearance, the fan blades of an open rotor engine resemble the propeller blades of a conventional turboprop engine. 
         [0005]    The use of contra-rotating rotor assemblies provides technical challenges in transmitting power from the power turbine to drive the blades of the respective two rotor assemblies in opposing directions. 
         [0006]    It would be desirable to provide an open rotor propulsion system utilizing a single rotating propeller assembly analogous to a traditional bypass fan which reduces the complexity of the design, yet yields a level of propulsive efficiency comparable to contra-rotating propulsion designs with a significant weight and length reduction. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0007]    An unducted thrust producing system has a rotating element with an axis of rotation and a stationary element. The rotating element includes a plurality of blades, and the stationary element has a plurality of vanes configured to impart a change in tangential velocity of the working fluid opposite to that imparted by the rotating element acted upon by the rotating element. The system includes an inlet forward of the rotating element and the stationary element. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0008]    The accompanying drawings, which are incorporated in and constitute a part of the specification, illustrate one or more embodiments and, together with the description, explain these embodiments. In the drawings: 
           [0009]      FIG. 1  is a cross-sectional schematic illustration of an exemplary embodiment of an unducted thrust producing system; 
           [0010]      FIG. 2  is an illustration of an alternative embodiment of an exemplary vane assembly for an unducted thrust producing system; 
           [0011]      FIG. 3  is a partial cross-sectional schematic illustration of an exemplary embodiment of an unducted thrust producing system depicting an exemplary compound gearbox configuration; 
           [0012]      FIG. 4  is a partial cross-sectional schematic illustration of an exemplary embodiment of an unducted thrust producing system depicting another exemplary gearbox configuration; 
           [0013]      FIG. 5  is a cross-sectional schematic illustration of another exemplary embodiment of an unducted thrust producing system; 
           [0014]      FIG. 6  is a cross-sectional schematic illustration of another exemplary embodiment of an unducted thrust producing system; 
           [0015]      FIG. 7  is a cross-sectional schematic illustration of another exemplary embodiment of an unducted thrust producing system; 
           [0016]      FIG. 8  is a cross-sectional schematic illustration of another exemplary embodiment of an unducted thrust producing system; 
           [0017]      FIG. 9  is a cross-sectional schematic illustration of another exemplary embodiment of an unducted thrust producing system; 
           [0018]      FIG. 10  is a cross-sectional schematic illustration of another exemplary embodiment of an unducted thrust producing system; 
           [0019]      FIG. 11  is a cross-sectional schematic illustration of another exemplary embodiment of an unducted thrust producing system; 
           [0020]      FIG. 12  is a cross-sectional schematic illustration of another exemplary embodiment of an unducted thrust producing system; 
           [0021]      FIG. 13  is a cross-sectional schematic illustration of another exemplary embodiment of an unducted thrust producing system; 
           [0022]      FIG. 14  is a cross-sectional schematic illustration of another exemplary embodiment of an unducted thrust producing system; and 
           [0023]      FIG. 15  is a cross-sectional schematic illustration taken along lines  15 - 15  of  FIG. 14  illustrating the inlet configuration of the unducted thrust producing system of  FIG. 14 . 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0024]    In all of the Figures which follow, like reference numerals are utilized to refer to like elements throughout the various embodiments depicted in the Figures. 
         [0025]      FIG. 1  shows an elevational cross-sectional view of an exemplary embodiment of an unducted thrust producing system  10 . As is seen from  FIG. 1 , the unducted thrust producing system  10  takes the form of an open rotor propulsion system and has a rotating element  20  depicted as a propeller assembly which includes an array of airfoil blades  21  around a central longitudinal axis  11  of the unducted thrust producing system  10 . Blades  21  are arranged in typically equally spaced relation around the centreline  11 , and each blade  21  has a root  23  and a tip  24  and a span defined therebetween. Unducted thrust producing system  10  includes a gas turbine engine having a gas generator  40  and a low pressure turbine  50 . Left- or right-handed engine configurations can be achieved by mirroring the airfoils of  21 ,  31 , and  50 . As an alternative, an optional reversing gearbox  55  (located in or behind the low pressure turbine  50  as shown in  FIGS. 3 and 4  or combined or associated with power gearbox  60  as shown in  FIG. 3 ) permits a common gas generator and low pressure turbine to be used to rotate the fan blades either clockwise or counterclockwise, i.e., to provide either left- or right-handed configurations, as desired, such as to provide a pair of oppositely-rotating engine assemblies as may be desired for certain aircraft installations. Unducted thrust producing system  10  in the embodiment shown in  FIG. 1  also includes an integral drive (power gearbox)  60  which may include a gearset for decreasing the rotational speed of the propeller assembly relative to the low pressure turbine  50 . 
         [0026]    Unducted thrust producing system  10  also includes in the exemplary embodiment a non-rotating stationary element  30  which includes an array of vanes  31  also disposed around central axis  11 , and each blade  31  has a root  33  and a tip  34  and a span defined therebetween. These vanes may be arranged such that they are not all equidistant from the rotating assembly, and may optionally include an annular shroud or duct  100  distally from axis  11  (as shown in  FIG. 2 ) or may be unshrouded. These vanes are mounted to a stationary frame and do not rotate relative to the central axis  11 , but may include a mechanism for adjusting their orientation relative to their axis  90  and/or relative to the blades  21 . For reference purposes,  FIG. 1  also depicts a Forward direction denoted with arrow F, which in turn defines the forward and aft portions of the system. As shown in  FIG. 1 , the rotating element  20  is located forward of the gas generator  40  in a “puller” configuration, and the exhaust  80  is located aft of the stationary element  30 . 
         [0027]    In addition to the noise reduction benefit, the duct  100  shown in  FIG. 2  provides a benefit for vibratory response and structural integrity of the stationary vanes  31  by coupling them into an assembly forming an annular ring or one or more circumferential sectors, i.e., segments forming portions of an annular ring linking two or more vanes  31  such as pairs forming doublets. The duct  100  may allow the pitch of the vanes to be varied as desired. 
         [0028]    A significant, perhaps even dominant, portion of the noise generated by the disclosed fan concept is associated with the interaction between wakes and turbulent flow generated by the upstream blade-row and its acceleration and impingement on the downstream blade-row surfaces. By introducing a partial duct acting as a shroud over the stationary vanes, the noise generated at the vane surface can be shielded to effectively create a shadow zone in the far field thereby reducing overall annoyance. As the duct is increased in axial length, the efficiency of acoustic radiation through the duct is further affected by the phenomenon of acoustic cut-off, which can be employed, as it is for conventional aircraft engines, to limit the sound radiating into the far-field. Furthermore, the introduction of the shroud allows for the opportunity to integrate acoustic treatment as it is currently done for conventional aircraft engines to attenuate sound as it reflects or otherwise interacts with the liner. By introducing acoustically treated surfaces on both the interior side of the shroud and the hub surfaces upstream and downstream of the stationary vanes, multiple reflections of acoustic waves emanating from the stationary vanes can be substantially attenuated. 
         [0029]    In operation, the rotating blades  21  are driven by the low pressure turbine via gearbox  60  such that they rotate around the axis  11  and generate thrust to propel the unducted thrust producing system  10 , and hence an aircraft to which it is associated, in the forward direction F. 
         [0030]    It may be desirable that either or both of the sets of blades  21  and  31  incorporate a pitch change mechanism such that the blades can be rotated with respect to an axis of pitch rotation either independently or in conjunction with one another. Such pitch change can be utilized to vary thrust and/or swirl effects under various operating conditions, including to provide a thrust reversing feature which may be useful in certain operating conditions such as upon landing an aircraft. 
         [0031]    Blades  31  are sized, shaped, and configured to impart a counteracting swirl to the fluid so that in a downstream direction aft of both rows of blades the fluid has a greatly reduced degree of swirl, which translates to an increased level of induced efficiency. Blades  31  may have a shorter span than blades  21 , as shown in  FIG. 1 , for example, 50% of the span of blades  21 , or may have longer span or the same span as blades  21  as desired. Vanes  31  may be attached to an aircraft structure associated with the propulsion system, as shown in  FIG. 1 , or another aircraft structure such as a wing, pylon, or fuselage. Vanes  31  of the stationary element may be fewer or greater in number than, or the same in number as, the number of blades  21  of the rotating element and typically greater than two, or greater than four, in number. 
         [0032]    In the embodiment shown in  FIG. 1 , an annular  360  degree inlet  70  is located between the fan blade assembly  20  and the fixed or stationary blade assembly  30 , and provides a path for incoming atmospheric air to enter the gas generator  40  radially inwardly of the stationary element  30 . Such a location may be advantageous for a variety of reasons, including management of icing performance as well as protecting the inlet  70  from various objects and materials as may be encountered in operation. 
         [0033]      FIG. 5  illustrates another exemplary embodiment of a gas turbine engine  10 , differing from the embodiment of  FIG. 1  in the location of the inlet  71  forward of both the rotating element  20  and the stationary element  30  and radially inwardly of the rotating element  20 . 
         [0034]      FIGS. 1 and 5  both illustrate what may be termed a “puller” configuration where the thrust-generating rotating element  20  is located forward of the gas generator  40 .  FIG. 6  on the other hand illustrates what may be termed a “pusher” configuration embodiment where the gas generator  40  is located forward of the rotating element  20 . As with the embodiment of  FIG. 5 , the inlet  71  is located forward of both the rotating element  20  and the stationary element  30  and radially inwardly of the rotating element  20 . The exhaust  80  is located inwardly of and aft of both the rotating element  20  and the stationary element  30 . The system depicted in  FIG. 6  also illustrates a configuration in which the stationary element  30  is located forward of the rotating element  20 . 
         [0035]    The selection of “puller” or “pusher” configurations may be made in concert with the selection of mounting orientations with respect to the airframe of the intended aircraft application, and some may be structurally or operationally advantageous depending upon whether the mounting location and orientation are wing-mounted, fuselage-mounted, or tail-mounted configurations. 
         [0036]      FIGS. 7 and 8  illustrate “pusher” embodiments similar to  FIG. 6  but wherein the exhaust  80  is located between the stationary element  30  and the rotating element  20 . While in both of these embodiments the rotating element  20  is located aft of the stationary element  30 ,  FIGS. 7 and 8  differ from one another in that the rotating element  20  of  FIG. 7  incorporates comparatively longer blades than the embodiment of  FIG. 8 , such that the root  23  of the blades of  FIG. 7  is recessed below the airstream trailing aft from the stationary element  30  and the exhaust from the gas generator  40  is directed toward the leading edges of the rotating element  20 . In the embodiment of  FIG. 8 , the rotating element  20  is more nearly comparable in length to the stationary element  30  and the exhaust  80  is directed more radially outwardly between the rotating element  20  and the stationary element  30 . 
         [0037]      FIGS. 9 ,  10 , and  11  depict other exemplary “pusher” configuration embodiments wherein the rotating element  20  is located forward of the stationary element  30 , but both elements are aft of the gas generator  40 . In the embodiment of  FIG. 9 , the exhaust  80  is located aft of both the rotating element  20  and the stationary element  30 . In the embodiment of  FIG. 10 , the exhaust  80  is located forward of both the rotating element  20  and the stationary element  30 . Finally, in the embodiment of  FIG. 11 , the exhaust  80  is located between the rotating element  20  and the stationary element  30 . 
         [0038]      FIGS. 12 and 13  show different arrangements of the gas generator  40 , the low pressure turbine  50  and the rotating element  20 . In  FIG. 12 , the rotating element  20  and the booster  300  are driven by the low pressure turbine  50  directly coupled with the booster  300  and connected to the rotating element  20  via the speed reduction device  60 . The high pressure compressor  301  is driven directly by the high pressure turbine  302 . In  FIG. 13  the rotating element  20  is driven by the low pressure turbine  50  via the speed reduction device  60 , the booster  303  is driven directly by the intermediate pressure turbine  306 , and the high pressure compressor  304  is driven by the high pressure turbine  305 . 
         [0039]      FIG. 15  is a cross-sectional schematic illustration taken along lines  15 - 15  of  FIG. 14  illustrating the inlet configuration of the unducted thrust producing system of  FIG. 14  as a non-axisymmetric, non-annular inlet. In the configuration shown, the inlet  70  takes the form of a pair of radially-opposed inlets  72  each feeding into the core. 
         [0040]    The gas turbine or internal combustion engine used as a power source may employ an inter-cooling element in the compression process. Similarly, the gas turbine engine may employ a recuperation device downstream of the power turbine. 
         [0041]    In various embodiments, the source of power to drive the rotating element  20  may be a gas turbine engine fuelled by jet fuel or liquid natural gas, an electric motor, an internal combustion engine, or any other suitable source of torque and power and may be located in proximity to the rotating element  20  or may be remotely located with a suitably configured transmission such as a distributed power module system. 
         [0042]    In addition to configurations suited for use with a conventional aircraft platform intended for horizontal flight, the technology described herein could also be employed for helicopter and tilt rotor applications and other lifting devices, as well as hovering devices. 
         [0043]    It may be desirable to utilize the technologies described herein in combination with those described in the co-pending applications listed above. 
         [0044]    The foregoing description of the embodiments of the invention is provided for illustrative purposes only and is not intended to limit the scope of the invention as defined in the appended claims. 
         [0045]    This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.