Abstract:
An aircraft propulsion system has a propulsive rotor assembly rotatable about an axis of rotation and comprising a plurality of blades and a rotationally fixed vane assembly located adjacent to the propulsive rotor assembly and arranged circumferentially around the axis of rotation. As airflow enters the propulsive rotor assembly, a portion of the airflow passes over the vane assembly which is configured to direct the airflow away from the rotor blades so as to reduce the relative velocity of the redirected airflow over the rotor blades. This results in a reduced tendency of the airflow through the propulsive rotor assembly to become choked.

Description:
[0001]    This invention claims the benefit of UK Patent Application No. 1114380.7, filed on 22 Aug. 2011, which is hereby incorporated herein in its entirety. 
       FIELD OF THE INVENTION  
       [0002]    This invention relates to aircraft propulsion systems and particularly, but not exclusively, to aircraft propulsion systems which utilise propfan or unducted fan type engines generally characterised by two contra-rotating propellers or open rotors. 
       BACKGROUND TO THE INVENTION  
       [0003]    It is known that increasing the bypass ratio of a conventional turbofan engine can reduce its fuel consumption and consequent level of CO 2  emissions. This characteristic has been exploited by engine manufacturers by progressively increasing the bypass ratios of modern turbofan engines. 
         [0004]    However, there is a limit to how much the bypass ratio can be increased as eventually the weight and drag penalties associated with the size of the required engine nacelle outweigh the reduction in fuel consumption. 
         [0005]    An alternative to the high bypass turbofan engine is the open rotor or unducted fan engine where the rotor or fan is not contained within the nacelle. This enables the fan or propeller to be larger in diameter which increases the bypass ratio while at the same time removing the need for a heavy drag-inducing nacelle. 
         [0006]    This in turn allows the open rotor engine to burn significantly less fuel (up to 30% in some instances) and offer associated reductions in emissions when compared to a conventional turbofan engine. 
         [0007]    An open rotor engine, although similar to a turboprop engine, is designed to operate efficiently at higher cruise velocities than turboprop engines. The primary difference between turboprop and open rotor engines is that the propeller blades on an open rotor engine have a higher solidity (generally by virtue of the greater number of propeller blades) than those of a turboprop engine. In addition, in contrast to a turboprop engine, an open rotor engine generates a portion of its total thrust from the engine&#39;s core exhaust. 
         [0008]    A problem with open rotor engines is that they generate higher noise levels than conventional turbofan engines, in which noise from the fan is muffled by the nacelle. 
         [0009]    Noise is of particular concern in a preferred open rotor arrangement which comprises two contra-rotating rotor assemblies. The wakes produced by the first (upstream) rotor are ‘chopped’ through by the rear (downstream) rotor. The intensity of the noise emitted by the wake interaction between the front rotor and the rear rotor is proportional to the strength of the wakes produced by the front rotor blades. The strength of the wakes produced by the blades of the front rotor can be reduced by decreasing the loading, or lift, generated by each individual blade. This can be accomplished without compromising the thrust produced by the engine by increasing the number of front rotor blades, i.e. each rotor blade is required to produce less lift in order to produce the same total engine thrust. 
         [0010]    In an open rotor engine arrangement the front rotor is subjected to axial Mach numbers equivalent to the forward flight speed of the aircraft. The high inlet velocity combined with the rotational speed of the rotor can result in the air entering the passage between two adjacent rotor blades having an even higher relative Mach number (approximately Mach 0.8). This may result in the flow regime through the open rotor suffering from choking. 
         [0011]    In a conventional turbo-fan engine the nacelle and intake to the fan diffuses the flow such that the inlet flow Mach number is lower than the flight speed of the aircraft when in a cruise condition but is higher than the flight speed of the aircraft at take-off or landing. Consequently the problem of choking of fan blades generally does not occur in such engines. 
         [0012]    The term choke margin is often used to describe the range of flow conditions relative to the choke point of the rotor. This can be defined as follows: 
         [0000]    
       
         
           
             
               
                 
                   
                     Choke 
                      
                     
                         
                     
                      
                     Margin 
                   
                   = 
                   
                     ( 
                     
                       
                         
                           Choke 
                            
                           
                               
                           
                            
                           Flow 
                         
                         - 
                         
                           Operating 
                            
                           
                               
                           
                            
                           Point 
                            
                           
                             
                                 
                             
                              
                             
                                 
                             
                           
                            
                           Flow 
                         
                       
                       
                         Choke 
                          
                         
                             
                         
                          
                         Flow 
                       
                     
                     ) 
                   
                 
               
               
                 
                   Eqn 
                    
                   
                       
                   
                    
                   
                     ( 
                     1 
                     ) 
                   
                 
               
             
           
         
       
     
         [0013]    An additional constraint on the rotor blades is that they must have a minimum thickness in order to provide the required structural strength to satisfy the bird strike requirement. This required minimum thickness puts an upper limit on the size of the area (throat area) between two adjacent blades. 
         [0014]    The high relative inlet Mach numbers and the limitation in throat area driven by the required minimum blade thickness can result in the airflow through the rotor becoming choked. Choking is particularly prevalent over the inboard portion of the blade since in this region the throat area is smallest and the blade is the thickest. Consequently, in order to avoid the flow becoming choked, it is often necessary to limit the number of blades on the rotor. However, in order to reduce the rotor noise it is desirable to increase the number of blades on the rotor. 
       STATEMENTS OF INVENTION  
       [0015]    According to a first aspect of the present invention there is provided an aircraft propulsion system comprising a first propulsive rotor assembly rotatable about an axis of rotation and comprising a plurality of first rotor blades, and a first vane assembly rotationally fixed about the axis of rotation, located adjacent to the first rotor assembly and arranged circumferentially around the axis of rotation, the first vane assembly being configured to direct an airflow entering the first propulsive rotor assembly so as to reduce the velocity of the redirected airflow relative to the first rotor blades. 
         [0016]    In an embodiment of the invention, the vane assembly is configured to turn the airflow such that it is more aligned with each of the rotor blades of the first rotor assembly. This decreases the magnitude and changes the direction of the relative velocity vector of the airflow entering the first rotor assembly. 
         [0017]    The reduction in the relative velocity of the airflow over the blades of the first rotor assembly improves the choke margin of the rotor by reducing the relative Mach number of the airflow entering the passage between two adjacent rotor blades. This enables the blade count of the rotor to be increased which improves the overall efficiency of the system whilst also reducing the level of noise emitted. 
         [0018]    The change in direction of the relative velocity vector also allows untwisting of the inboard portion of the rotor. In other words, the aerofoil section of the blade can be more closely aligned with the axis of rotation of the rotor assembly. This results in an increase in the throat area between two adjacent rotor blades which further improves the choke margin. 
         [0019]    Optionally, the aircraft propulsion system further comprises a second propulsive rotor assembly comprising a plurality of second rotor blades and positioned on an opposite side of the first propulsive rotor assembly to the first vane assembly, and rotatable about the axis of rotation. 
         [0020]    Optionally, the first propulsive rotor assembly and the second propulsive rotor assembly rotate in opposite directions to one another. 
         [0021]    Optionally, the first vane assembly comprises a plurality of vanes and at least one of the plurality of vanes is formed as an aerofoil. 
         [0022]    By forming the vanes as aerofoils, the oncoming airflow can be more readily redirected to change the direction and magnitude of the relative velocity vector. 
         [0023]    In an embodiment of the invention, the span of the vanes is less than the span of the rotor blades. 
         [0024]    In other embodiments of the invention the vanes may have a span which is substantially the same as that of the rotor blades. 
         [0025]    In another embodiment of the invention the vanes may be twisted along their span. 
         [0026]    Optionally, the plurality of vanes is arranged asymmetrically around the axis of rotation. 
         [0027]    By arranging the vanes asymmetrically around the axis of rotation it is possible to mitigate engine installation effects on the rotor such as, for example, a wake propagating from an engine mounting pylon. 
         [0028]    Optionally, each of the plurality of vanes has at least one of a span, chord, camber, circumferential lean or sweep which differs from the corresponding span, chord, camber, circumferential lean or sweep of an adjacent vane. 
         [0029]    When an open rotor fan engine is mounted to an aircraft, the flow field around the engine assembly will generally be asymmetric as a consequence of, for example, the influence of the engine mounting pylon. 
         [0030]    By altering the geometry of one or more of the vanes it becomes possible to asymmetrically modify the airflow entering the rotor assembly to compensate for the influence of the engine pylon. In addition to providing an improvement in choke margin, such an axisymmetric vane arrangement also provides a reduction in noise by reducing the need for the blades to pass through a cyclically varying flow field. 
         [0031]    Optionally, one or more of the plurality of vanes comprises a winglet formed on a radially outermost portion of the one or more vanes. 
         [0032]    The use of a winglet formed at a tip of the vane aids in decreasing the strength of tip vortices from the vane. The reduction in vane tip vortex strength reduces the noise caused by rotor interactions with the vortex. By limiting the effect of the vortex, the airflow into the rotor is smoothed which can in turn improve the choke margin and thus the aerodynamic efficiency of the rotor assembly. 
         [0033]    Optionally, one or more of the plurality of vanes comprises a respective one or more variable pitch vanes. 
         [0034]    The ability to vary the pitch of the vanes enables the degree to which the airflow across the rotor is modified to be varied depending on the operating conditions of the propulsion system. 
         [0035]    The pitch of the vanes is therefore adjusted in concert with the pitch of the rotor blades in order to cope with the changes in flight speed. For example, since choking is more of a problem at cruise conditions due to the higher flight Mach number, this portion of the flight envelope may require greater pitch angle adjustment of the rotors and vanes. 
         [0036]    Optionally, the aircraft propulsion system further comprises a vane control module operable to collectively alter a pitch of each of the one or more variable pitch vanes. 
         [0037]    Optionally, the aircraft propulsion system further comprises a housing, wherein at least one of the plurality of vanes is movable between a first, stowed position in which the at least one vane is accommodated at least partially within the housing, and a second deployed, position in which the at least one vane protrudes from the housing and extends into the airflow entering the first propulsive rotor assembly. 
         [0038]    By retracting the vanes when they are not required, it is possible to achieve a reduction in aerodynamic drag which makes the propulsion system more efficient. 
         [0039]    Optionally, the aircraft propulsion system further comprises a second vane assembly rotationally fixed about the axis of rotation, arranged circumferentially around the axis of rotation, and positioned between the first and second propulsive rotor assemblies, the second vane assembly being configured to direct an airflow exiting the first propulsive rotor assembly so as to reduce the velocity of the redirected airflow relative to the surface of the second rotor blades. 
         [0040]    A second vane assembly may be mounted on the nacelle between the front and rear rotors and arranged circumferentially around the axis of rotation. The second vane assembly receives the highly swirled flow exiting the first rotor assembly and removes a portion of this swirl such that the relative velocity (and hence relative Mach number) of the airflow entering the second rotor assembly is reduced. The second vane assembly may be implemented with or without the first vane assembly. 
         [0041]    According to a second aspect of the present invention there is provided a method of controlling an aircraft propulsion system comprising a first propulsive rotor assembly rotatable about an axis of rotation and comprising a plurality of first rotor blades, and a plurality of vanes located adjacent to the first propulsive rotor assembly and arranged circumferentially around the axis of rotation, the plurality of vanes being configured to direct an airflow entering the first propulsive rotor assembly so as to reduce the velocity of the redirected airflow relative to the first rotor blades, the method comprising the step of:
       collectively changing a pitch of the plurality of vanes.       
 
         [0043]    Optionally, the aircraft propulsion system further comprises a housing, the plurality of vanes being accommodated at least partially within the housing, the method comprising the additional initial step of:
       moving each of the plurality of vanes from a first, stowed position to a second, deployed position in which each vane extends from the housing into an airflow entering the first propulsive rotor assembly.       
 
         [0045]    Optionally the method comprises the step of radially extending each of the plurality of vanes from a stowed position to a deployed position. 
         [0046]    Optionally the method comprises the step of rotating each of the plurality of vanes from a stowed position to a deployed position. 
         [0047]    According to a third aspect of the present invention there is provided a nacelle comprising an aircraft propulsion system including a first propulsive rotor assembly rotatable about an axis of rotation and a rotationally fixed first vane assembly positioned adjacent to the first propulsive rotor assembly and arranged circumferentially around the axis of rotation, wherein an axis of the nacelle is aligned with the axis of rotation. 
         [0048]    According to a fourth aspect of the invention, there is provided method of reducing the noise emitted by an aircraft propulsion system comprising at least one propulsive rotor assembly having a plurality of rotor blades and rotatable about an axis of rotation, the method comprising the steps of installing a plurality of vanes in a circumferential arrangement around the axis of rotation to thereby reduce the velocity of an airflow over the rotor blades and to thereby reduce the rotor inlet Mach number and modifying the propulsive rotor blade assembly by adding at least one additional rotor blade to thereby reduce the noise generated by the rotor assembly when in use. 
         [0049]    Other aspects of the invention provide devices, methods and systems which include and/or implement some or all of the actions described herein. The illustrative aspects of the invention are designed to solve one or more of the problems herein described and/or one or more other problems not discussed. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0050]    There now follows a description of an embodiment of the invention, by way of non-limiting example, with reference being made to the accompanying drawings in which: 
           [0051]      FIG. 1  is a schematic cross-sectional view of a conventional open-rotor turbofan aircraft engine; 
           [0052]      FIG. 2  is a diagrammatic representation of the airflow over the open-rotor blades of the prior art engine of  FIG. 1 ; 
           [0053]      FIG. 3  is a schematic cross-sectional view of an open-rotor turbofan aircraft engine according to a first embodiment of the invention; 
           [0054]      FIG. 4   a  is a partial side view of the engine of  FIG. 3  showing a vane and a rotor blade; 
           [0055]      FIG. 4   b  is a plan view of a vane of the embodiment of  FIG. 3 ; 
           [0056]      FIG. 5  is a diagrammatic representation of the airflow over the open-rotor blades of the engine of  FIG. 3 ; 
           [0057]      FIGS. 6   a  and  6   b  are schematic elevational views of an engine nacelle showing symmetric and asymmetric arrangements respectively of a vane assembly according to the embodiment of  FIG. 3 ; 
           [0058]      FIG. 6   c  is a schematic elevational view of an aircraft having pylon mounted nacelles with asymmetric arrangements of vane assemblies according to the embodiment of  FIG. 3 ; 
           [0059]      FIG. 7  is a perspective schematic view of a vane of the embodiment of  FIG. 3 ; 
           [0060]      FIG. 8  is a schematic cross-sectional view of an open-rotor turbofan aircraft engine according to a second embodiment of the invention; 
           [0061]      FIG. 9  is a schematic, partial, cross-sectional view of the engine of  FIG. 8  having a radially deployable vane assembly and showing one vane in a partially radially retracted position; 
           [0062]      FIG. 10   a  is a schematic elevational view of a portion of an engine nacelle in which the vane assembly is rotatably retractable and showing a vane in the retracted position; 
           [0063]      FIG. 10   b  shows the vane assembly of  FIG. 10   a  in an extended position; and 
           [0064]      FIG. 11  is a schematic, partial, cross-sectional view of the engine of  FIG. 8  having a rotatably retractable vane assembly and showing nacelle shaping by a retracted vane. 
       
    
    
       [0065]    It is noted that the drawings may not be to scale. The drawings are intended to depict only typical aspects of the invention, and therefore should not be considered as limiting the scope of the invention. In the drawings, like numbering represents like elements between the drawings. 
       DETAILED DESCRIPTION 
       [0066]    Referring to  FIG. 1 , a conventional open-rotor turbofan engine assembly is designated generally by the reference numeral  10 . 
         [0067]    The open rotor engine assembly  10  comprises an engine core  12 , a nacelle  14  which separates the core airflow  16  from an external airflow  18 , a first rotor assembly  20  comprising a plurality of first rotor blades  22 , and a second rotor assembly  80  comprising a plurality of second rotor blades  82 . 
         [0068]    Both the first and second rotor assemblies  20 ,  80  rotate about an axis of rotation  24 . 
         [0069]    For a given flight condition, the velocity of the external airflow  18  is sufficient to cause a choked region  26  to form over a portion of the span of the first rotor  20 , as indicated in  FIG. 1 . 
         [0070]      FIG. 2  shows, in diagrammatic form, the airflow  18  entering the first rotor assembly  20  of the engine assembly  10 . A velocity triangle  30  represents the vector components of the airflow  18  at a given rotor radius  28 . 
         [0071]    A rotational velocity vector  32 , which represents the circumferential rotor velocity, is added to the axial velocity vector  34 , representing the axial flight speed of the aircraft, to create a relative velocity vector  36  which is oriented at an angle  38  to the axial velocity vector  34 . 
         [0072]    The relative velocity vector  36  is greater in magnitude than the axial velocity vector  34 . This means that the velocity of the airflow in the plane of the rotor blade section is greater than the velocity of the airflow along the axis of the rotor. 
         [0073]    The airflow velocity can be converted to Mach number by the well known relationship: 
         [0000]    
       
         
           
             
               
                 
                   M 
                   = 
                   
                     v 
                     
                       γRT 
                     
                   
                 
               
               
                 
                   Eqn 
                    
                   
                       
                   
                    
                   
                     ( 
                     2 
                     ) 
                   
                 
               
             
           
         
       
       
         
           
             where: M=Mach number;
           v=airflow velocity;   γ=specific heat ratio;   R=universal gas constant; and   T=absolute temperature.   
         
           
         
       
     
         [0079]    Each of the first rotor blades  22  has an aerofoil-shaped cross sectional profile  23  having a maximum thickness  42 , and being substantially aligned with the relative velocity vector  36 . The angle of alignment of the first rotor blade sections  23  is termed the stagger angle  40 . 
         [0080]    The maximum thickness  42  is determined largely by a structural requirement for the rotor blade  22  to be capable of withstanding a bird strike while in flight without comprising the operational integrity of the engine assembly  10 . 
         [0081]    The relative spacing of the first rotor blades  22  creates a passage  44  between adjacent front rotor blades  22 , having an inlet area  46  which narrows to a smaller minimum area defined as a throat area  48 . The choice of aerofoil profile shape in combination with the thickness  42  and the stagger  40  dictates the size of the throat area  48 . 
         [0082]    The ratio of the inlet area  46  to the throat area  48  for the choking condition to occur at a given Mach number is governed by the known compressible flow relationship: 
         [0000]    
       
         
           
             
               
                 
                   
                     A 
                     
                       A 
                       * 
                     
                   
                   = 
                   
                     
                       1 
                       M 
                     
                      
                     
                       { 
                       
                         
                           2 
                           
                             ( 
                             
                               γ 
                               + 
                               1 
                             
                             ) 
                           
                         
                          
                         
                           ( 
                           
                             1 
                             + 
                             
                               
                                 ( 
                                 
                                   
                                     γ 
                                     - 
                                     1 
                                   
                                   2 
                                 
                                 ) 
                               
                                
                               
                                 M 
                                 2 
                               
                             
                           
                           ) 
                         
                       
                       } 
                     
                   
                 
               
               
                 
                   Eqn 
                    
                   
                       
                   
                    
                   
                     ( 
                     3 
                     ) 
                   
                 
               
             
           
         
       
       
         
           
             where: A=inlet area;
           A*=throat area at which choking will occur;   M=relative Mach number at the inlet  46  to the rotor; and   γ=specific heat ratio.   
         
           
         
       
     
         [0087]    If the ratio of the inlet area  46  to the throat area  48  for the first rotor  20  is greater than NA* for the relative inlet Mach number at a flight condition, i.e. 
         [0000]    
       
         
           
             
               
                 ( 
                 
                   
                     Inlet 
                      
                     
                         
                     
                      
                     Area 
                   
                   
                     Throat 
                      
                     
                         
                     
                      
                     Area 
                   
                 
                 ) 
               
               &gt; 
               
                 A 
                 
                   A 
                   * 
                 
               
             
             , 
           
         
       
     
         [0000]    choking will occur. This will result in a loss in efficiency, an increase in drag, and limit to the thrust which the engine can provide. Thus choking is governed by the relative Mach number and the area ratio, 
         [0000]    
       
         
           
             
               ( 
               
                 
                   Inlet 
                    
                   
                       
                   
                    
                   Area 
                 
                 
                   Throat 
                    
                   
                       
                   
                    
                   Area 
                 
               
               ) 
             
             . 
           
         
       
     
         [0088]    Referring to  FIG. 3 , an open rotor engine assembly according to an embodiment of the invention is designated generally by the reference numeral  100 . Features of the engine assembly  100  which correspond to those of engine assembly  10  have been given corresponding reference numerals for ease of reference. 
         [0089]    The engine assembly  100  includes all the features of engine assembly  10  with the addition of a first vane assembly  60  comprising a plurality of first vanes  62 . The first vane assembly  60  is mounted on the nacelle  14  and is circumferentially arranged around the axis of rotation  24 . 
         [0090]    As shown in  FIG. 4   a , the first vanes  62  have a span  63  which extends slightly beyond the region  26  for which choking or choke margin is a concern. The cross sectional profile  64  of a first vane  62  can be suitably designed by using velocity diagrams (see  FIG. 5 ) to achieve a desirable combination of rotor blade inlet Mach number and area ratio at any given point in the span  63 . 
         [0091]    The tip cross section  66  (see  FIG. 4   b ) of a first vane  62  which extends beyond the choked region  26  of the first rotor  20  can be designed with zero camber in order to produce very little or no lift at the tip of the vane. Designing the tip region  65  of the airfoil to carry zero lift will limit the effect of tip vortices known to occur at the free end of wings. Other known methods of diminishing the strength of tip vortices that may be generated by the first vanes  62  can be employed, such as winglets or elliptical planform shaping. 
         [0092]      FIG. 5  shows in diagrammatic form the airflow  18  through the first vanes  62  and first rotor assembly  20  of the engine assembly  100 . A velocity triangle  130  represents in vector form the components of the airflow  18  at the rotor radius  28 . 
         [0093]    The free stream airflow  18 , represented by axial velocity vector  34 , entering the first vane assembly  60  is turned through a small angle  150  by the first vane  62 . As shown in  FIG. 5 , the first vane  62  has an aerofoil-shaped cross-sectional profile  64 . 
         [0094]    Having passed over the vane section  64  the airflow has a vane exit velocity vector  134 . The rotational velocity vector  32  of the first rotor  20  combines with the vane exit velocity vector  134  to create a new relative inlet velocity vector  136  which is oriented at an angle  138  to the axial velocity vector  34 . 
         [0095]    The new relative inlet velocity vector  136  is smaller in magnitude than the corresponding relative inlet velocity vector  36  for the engine  10  (i.e. without the first vane assembly  60 ). In addition, the relative inlet velocity vector  136  is more axially aligned with the axis of rotation  24  than is the corresponding relative inlet velocity vector  36  (i.e. angle  138  is smaller than angle  38 ). 
         [0096]    The rotor sections  23  of engine  100  are substantially aligned with the new relative inlet velocity vector  136  to create a new stagger angle  140 . 
         [0097]    This new stagger angle  140  is smaller than the corresponding stagger angle  40  in engine  10  which results in the passage  44  having a new inlet area  146  and throat area  148 . 
         [0098]    Since, as described above, the blade sections  23  of engine  100  are more axially aligned than the corresponding blade sections  23  of engine  10 , the throat area  148  can be larger than the throat angle  48  of engine  10 . This lowers the airflow velocity through the passage  44  which has the effect of decreasing the relative Mach number of the airflow. 
         [0099]    This decrease in relative Mach number together with the increase in throat area  148  results in an improvement in the choke margin. 
         [0100]    It is to be noted that the current invention can provide a performance benefit to the engine, for example by decreasing the ratio of 
         [0000]    
       
         
           
             
               ( 
               
                 A 
                 
                   A 
                   * 
                 
               
               ) 
             
              
             
                 
             
              
             or 
              
             
                 
             
              
             
               ( 
               
                 
                   Inlet 
                    
                   
                       
                   
                    
                   Area 
                 
                 
                   Throat 
                    
                   
                       
                   
                    
                   Area 
                 
               
               ) 
             
           
         
       
     
         [0000]    without eliminating choking in the blade passages. In this situation the value of the choke margin according to Equation (1) will be zero, i.e. the same as if the vanes were not present and the flow was choked. However the vanes will still decrease the Mach number relative to the blade passage which in turn will decrease the strength of the shock which forms at the passage throat. A weaker shock will produce less aerodynamic loss and the system performance will therefore be improved. 
         [0101]    In an alternative embodiment of the invention (see  FIGS. 6   a ,  6   b  and  6   c ), the individual first vanes  62  that are arranged circumferentially on the nacelle  14  could each have different geometries to one another in order to accommodate an asymmetric flow field around the engine nacelle  14 . 
         [0102]    This asymmetric flow field arises from the airflow passing over the aircraft structure  160  not being symmetrically disposed relative to the engine axis  24 . One major effect on the flow  18  experienced by an open rotor engine  10  is that caused by the engine mounting pylon  162  which creates a strong wake near the inlet plane of the front rotor assembly  20 . 
         [0103]    The velocity field in the airflow around the engine can be analysed, accounting for these installation effects, using modern computational methods which are known in the art. Based on such an analysis the span  63 , chord  67 , stack  68 , camber  69  and sweep  70  (see  FIG. 7 ) of individual first vanes  62  can be suitably chosen by one skilled in the art. 
         [0104]      FIG. 6   a  shows a schematic elevational view of an engine nacelle  14  showing an axisymmetric distribution of vanes  62 , each having a common geometry. 
         [0105]      FIG. 6   b  shows a similar view to that of  FIG. 6   a  but with a non-axisymmetric distribution of vanes  62  and also showing vanes  62  having a variety of geometries. This variation in geometry might include, for example, vanes having greater span  170 , lesser span  172 , varying stack (shown as circumferential lean  180 , 182 ), and the addition of winglets  190 , 192 . 
         [0106]      FIG. 6   c  shows a schematic cross-sectional view of an aircraft  160  having two open rotor engines  100  mounted to respective pylons  162 . The engines  100  incorporate a non-axisymmetric arrangement of first vanes  62  such as that shown in  FIG. 6   b  which are upstream of the first rotor assembly  20 .  FIG. 6   c  is representative of both pusher-type and puller-type open-rotor arrangements. 
         [0107]    In a further embodiment ( FIGS. 10   a ,  10   b  and  11 ) each of the first vanes  62  may be deployable. That is to say the vanes  62  can be retracted into the nacelle for those portions of the flight envelope for which they do not provide an aerodynamic benefit to the first rotor  20 , for example during low speed flight operation. The vanes  62  may then be extended for the portions of the flight envelope for which they are beneficial, such as, for example, flying at cruising altitude. 
         [0108]    In one arrangement, the first vanes  62  may be extended radially from a retracted position in which they are accommodated either entirely, or substantially entirely, within the outer surface of the nacelle  14 . Any suitable actuating mechanism (not shown) could be used to extend and retract the vanes  62  such as, for example, hydraulic cylinders or electric motors. 
         [0109]    In an alternative arrangement (shown in  FIGS. 10   a  and  10   b ), the first vanes  62  may be hinged at a pivot  350  which is connected to the nacelle structure  351 . The vanes  62  may be retracted by pivoting the vanes  62  in a circumferential direction by means of an actuator  352 . In this arrangement, when the vane is in its retracted position, one of the vane surfaces  355  remains flush with the outer surface  353  of the nacelle  14 . 
         [0110]    In such an embodiment, the first vanes  62  may be attached to a nacelle filler  354  that is arranged to extend approximately perpendicularly to the vane  62  near the pivot axis  350 . As a result, when the vanes  62  are extended as shown in  FIG. 10   b , the nacelle filler  354  will rotate to be flush with the nacelle surface  353  and will fill the void left by the extended vane  62 . Other means such as, for example, a sliding panel might also be used to fill this void space. 
         [0111]    In a variation of this embodiment exemplified by  FIGS. 10   a  and  11 , a surface  355  of the rotatably deployable first vanes  62  protrudes above the outer surface  353  of the nacelle  14  when the vanes are in their retracted position. In this arrangement, the surface  355  of the vane  62  which is exposed to the inlet flow  18  is suitably shaped to provide an aerodynamic benefit (i.e. nacelle shaping) to the flow  18  entering the first rotor  20 . 
         [0112]    In yet another embodiment (not shown), the first vanes  62  can be arranged to have variable pitch. Variable pitch enables the vanes  62  to modify the inlet airflow  18  such that the resulting inlet flow velocity vector  136  is optimised for a range of operating conditions of the aircraft, such as, for example, take-off, climb and cruise. This variation in the pitch of the vanes  62  may or may not be linked to changes in the pitch angle of the rotor blades  22 . 
         [0113]    Referring to  FIG. 8 , an open rotor engine assembly according to a second embodiment of the invention is designated generally by the reference numeral  200 . Features of the engine assembly  200  which correspond to those of engine assembly  100  have been given corresponding reference numerals for ease of reference. 
         [0114]    The engine assembly  200  includes all the features of engine assembly  100  with the addition of a second vane assembly  90  comprising a plurality of second vanes  162 . The second vane assembly  90  is mounted on the nacelle  14  between the first and second rotor assemblies  20 , 80  and is circumferentially arranged around the axis of rotation  24 . 
         [0115]    The present invention may be embodied in other specific forms without departing from its spirit or essential characteristics. The described embodiments are to be considered in all respects only as illustrative and not restrictive. The scope of the invention is therefore indicated by the appended claims rather than by the foregoing description. All changes which come within the meaning and range of equivalency of the claims are to be embraced within their scope.