Abstract:
A variable geometry inlet system of an aircraft engine includes an inlet duct. The inlet duct includes at least first and second sections moveable between extended and retracted positions such that the inlet duct defines a variable axial length of an inlet passage for selective flight conditions. The inclusion of acoustic treatment may assist in controlling noise.

Description:
TECHNICAL FIELD 
       [0001]    The described subject matter relates generally to aeroengines, and more particularly to aircraft engines inlet systems. 
       BACKGROUND OF THE ART 
       [0002]    It is well understood that significant sources of noise generated by aircraft gas turbine engines are the result of viscous wake and rotor turbulence interaction. Such generated noise may propagate forwardly to thereby result in community noise particularly when the aircraft approaches an airport for landing. One of the conventional approaches toward suppressing noise generated in this manner, is to line the inlet with sound-absorbing panelling. However, because of the close proximity of the fan or compressor to the inlet frontal plane, and the lack of acoustic shielding in the forward direction, a significant portion of the noise may still propagate forward out of the inlet duct. 
         [0003]    Accordingly, there is a need to provide an improved engine inlet system for aircraft gas turbine engines. 
       SUMMARY 
       [0004]    In one aspect, there is provided a variable geometry inlet system of an aeroengine comprising an inlet duct for directing an air flow from an opening in the inlet duct to a compressor, the inlet duct including at least first and second sections, the first and the second sections movable relative to one another between an extended position and a retracted position, the first and second sections in the extended position defining an axially longer inlet duct than when in the retracted position, the inlet duct extending continuously from the opening to the compressor in both positions and providing only one inlet path for the air flow from the opening to the compressor regardless of the inlet duct position. 
         [0005]    In another aspect, there is provided a turbofan aeroengine comprising a fan and compressor section, a combustion section, a turbine section and a nacelle surrounding at least the fan and compressor section, the nacelle including a main portion in a stationary relationship with the engine, the main portion having an annular outer skin and an inner barrel radially spaced apart from each other, the nacelle including an inlet cowl having a lip portion connected with an annular outer skin and an annular inner wall radially spaced apart from each other, the inlet cowl being disposed upstream of the main portion with respect to an air flow entering the nacelle through a front opening defined by the lip portion of the inlet cowl, the inlet cowl being operatively connected to the main portion and translatable between a retracted position in which the outer skins of the respective inlet cowl and main portion of the nacelle are immediately axially adjacent each other and in which the annular inner wall of the inlet cowl and the inner barrel of the main portion are inserted one into the other, and an extended position in which the outer skins of the respective inlet cowl and main portion of the nacelle are positioned axially spaced apart from each other and in which the annular inner wall of the inlet cowl and the inner barrel of the main portion are inserted one into the other less than the insertion in the retracted position, the annular inner wall of the inlet cowl and the inner barrel of the main portion thereby forming a length-variable inlet duct for directing the air flow toward the fan and compressor section. 
         [0006]    In a further aspect, there is provided a turbofan aeroengine comprising a fan and compressor section, a combustion section, a turbine section and a nacelle surrounding at least the fan and compressor section, the nacelle including a main portion in a stationary relationship with the engine, the main portion having an annular outer skin and an inner barrel radially spaced apart from each other, the nacelle including an inlet cowl disposed upstream of the main portion with respect to an air flow entering the nacelle through a front opening defined by an annular lip portion of the inlet cowl, the lip portion being affixed to an annular outer skin of the inlet cowl, and a plurality of circumferentially distributed plates being pivotally connected about a tangentially pivoting axis at a respective front edge thereof to the lip portion to form an annular inner wall radially spaced apart from the outer skin of the inlet cowl, the annular inner wall being in an variable truncated conical profile having a fixed diameter at the front edge of the respective plates and a variable diameter at a rear edge of the respective plates, the rear edge of the respective plates being opposite to the front edge of the respective plates, the inlet cowl being operatively connected to the main portion and being axially translatable between a retracted position in which the outer skins of the respective inlet cowl and main portion of the nacelle are immediately axially adjacent each other with the inner barrel of the main portion being partially inserted into the inner wall of the inlet cowl, and an extended position in which the outer skins of the respective inlet cowl and the main portion of the nacelle are positioned axially spaced apart from each other with the inner barrel of the main portion being partially inserted into the inner wall of the inlet cowl less than the insertion in the retracted position, the annular inner wall of the inlet cowl and the inner barrel of the main portion thereby forming a length-variable inlet duct for directing the air flow. 
         [0007]    Further details of these and other aspects of the described subject matter will be apparent from the detailed description and drawings included below. 
     
    
     
       DESCRIPTION OF THE DRAWINGS 
         [0008]    Reference is now made to the accompanying figures in which: 
           [0009]      FIG. 1  is a schematic cross-sectional view of a turbofan aeroengine as an example illustrating the application of the described subject matter; 
           [0010]      FIG. 2  is a schematic partial cross-sectional view of a turbofan aeroengine having a variable geometry inlet system in a retracted position according to one embodiment; 
           [0011]      FIG. 3  is a schematic partial cross-sectional view of the turbofan aeroengine of  FIG. 2 , showing the variable geometry inlet system in an extended position; 
           [0012]      FIG. 4  is a schematic partial cross-sectional view of a turbofan aeroengine having a variable geometry inlet system in a retracted position according to another embodiment; 
           [0013]      FIG. 5  is a schematic partial cross-sectional view of the turbofan aeroengine of  FIG. 4  showing the variable geometry inlet system in an extended position; 
           [0014]      FIG. 6  is a schematic partial perspective view of the turbofan aeroengine having the variable geometry inlet system of  FIG. 4  in such a retracted position, with a portion of an outer skin of an inlet cowl cutaway; and 
           [0015]      FIG. 7  is a schematic partial perspective view of the turbofan aeroengine having the variable geometry inlet system in the extended position as shown in  FIG. 5 , with a portion of an outer skin of the inlet cowl cut away. 
       
    
    
       [0016]    It will be noted that throughout the appended drawings, like features are identified by like reference numerals. 
       DETAILED DESCRIPTION 
       [0017]      FIG. 1  illustrates a turbofan aeroengine as an example of the application of the described subject matter, which includes an outer bypass duct or nacelle  10 , a core casing  13 , a low pressure spool assembly seen generally at  12  which includes a fan assembly  14 , a low pressure compressor assembly  16  and a low pressure turbine assembly  18 , and a high pressure spool assembly seen generally at  20  which includes a high pressure compressor assembly  22  and a high pressure turbine assembly  24 . 
         [0018]    The core casing  13  surrounds the low and high pressure spool assemblies  12  and  20  in order to define a main fluid path (not numbered) therethrough including a combustor  26 . 
         [0019]    It should be noted that the terms axial, radial and circumferential are defined with respect to a main engine axis  28 . The terms downstream and upstream are defined with respect to the direction of an air flow indicated by arrow  30 , entering into and passing through the engine. 
         [0020]    Referring to  FIGS. 1-3 , the nacelle  10  of the aeroengine according to one embodiment, surrounds at least the fan assembly  14  and the low pressure and high pressure compressor assemblies  16 ,  22  and may be configured to provide a variable geometry inlet system  11   a  for the aeroengine. The nacelle  10  may include a main portion  32  in a stationary relationship with the engine, for example by being connected to the core casing  13  by a plurality of circumferentially spaced struts  34  (see  FIG. 1 ). The main portion  32  of the nacelle  10  may be connected by a support structure (not shown) to an aircraft (not shown). The main portion  32  may have an annular outer skin  36  and an inner barrel  38  radially spaced apart from each other. The inner barrel  38  may include a front section  40  extending forward out of a front edge  42  of the annular outer skin  36 , thereby positioning the front section  40  upstream of the front edge  42  of the annular outer skin  32 . 
         [0021]    The nacelle  10  may include an inlet cowl  44  which has an annular outer skin  48  and an annular inner wall  50  radially spaced apart from each other. The inlet cowl  44  may have a lip portion  46  at the upstream end thereof which is connected with the annular outer skin  48  and the inner wall  50 . The inlet cowl  44  may be disposed upstream of the main portion  32  of the nacelle  10  and may be operatively connected to the main portion  32 , for example by means of a track system (not shown) which is known in the art, such that the inlet cowl  44  is translatable between a retracted position as shown in  FIG. 2  in which the outer skins  48 ,  36  of the respective inlet cowl  44  and the main portion  32  of the nacelle  10  are positioned immediately axially adjacent each other while the annular inner wall  50  of the inlet cowl  44  and the inner barrel  38  of the main portion  32  are inserted one into the other, and an extended position as shown in  FIG. 3  in which the outer skins  48 ,  36  of the respective inlet cowl  44  and the main portion  32  of the nacelle  10  are positioned axially spaced apart from each other while the inner wall  50  of the inlet cowl  44  and the inner barrel  38  of the main portion  32  are inserted one into the other less than the insertion in the retracted position as shown in  FIG. 2 . Therefore, the annular inner wall  50  of the inlet cowl  44  and the inner barrel  38  (including its front section  40 ), in combination form a length-variable inlet duct. 
         [0022]    The variable geometry inlet duct has a front opening  52  defined by the annular lip portion  46  of the inlet cowl  44  for intake of the airflow  30 . The inlet duct directs the air flow  30  towards and to pass the fan rotor  14 . A front section (not numbered) of the inlet duct defined by the annular inner wall  50  of the inlet cowl  44  may be movable between the extended position and the retracted position with respect to a rear section (not numbered) of the inlet duct defined by the inner barrel  38  of the main portion  32 , to thereby define a variable axial length of the air inlet passage which is continuous from the front opening  52  to the fan rotor  14  without any secondary inlet opening being formed between the front and rear sections of the inlet duct, regardless of the position of the front section of the inlet duct (the position of the inlet cowl  44 ). 
         [0023]    The annular inner wall  50  of the inlet cowl  44  according to one embodiment, may be affixed to the lip portion  46  and may be slidingly inserted into a front section  40  of the inner barrel  38  of the main portion  32 , to form a telescoping configuration. The annular inner wall  50  of the inlet cowl  44  may define a front diameter corresponding to an inner diameter of the lip portion  46  of the inlet cowl  44  and may form a rear diameter thereof which may be greater than the front diameter of the annular inner wall  50  but slightly smaller than the inner diameter of the front section  40  of the inner barrel  38  of the main portion  32 . The front section  40  of the inner barrel  38  may have a substantially consistent diameter along the length thereof to thereby be substantially cylindrical. This telescoping configuration allows the annular wall  50  to be fully inserted into the front section  40  of the inner barrel  38  of the main portion  32 . In such a case, the annular front section  40  of the inner barrel  38  is fully received within an annular space between the outer skin  48  and the inner wall  50  of the inlet cowl  44  when the inlet cowl  44  is translated from the extended position to the retracted position. In this telescoping configuration the annular inner wall  50  of the inlet cowl  44  and the front section  40  of the inner barrel  38  may each be made of a respective metal ring of a single piece component. 
         [0024]    In the retracted position, the outer skin  48  of the inlet cowl  44  may be positioned immediately axially adjacent the front edge  42  of the annular outer skin  36  of main portion  32 , without a substantial axial gap therebetween, to thereby provide a low drag profile of the nacelle  10  for flight conditions such as cruise flight and take-off. In aircraft approach operations drag is less important for fuel consumption and therefore, the inlet cowl  44  can be translated to its extended position to increase the length/diameter ratio of the inlet duct formed within the nacelle  10 , which may help with reduction of noise levels propagated through the inlet duct. 
         [0025]    In order to further increase noise attenuation, the inner wall  50  of the inlet cowl  44  and the front section  40  of the inner barrel  38  of the main portion  32  may be provided with acoustic treatment capabilities, for example, by providing perforations therethrough or noise absorption material thereon to define a variable-geometry acoustic treatment area on the inner surface of the length-variable inlet duct formed within the nacelle  10 . When the inlet cowl  44  is in the extended position the acoustic treatment area defined by the inner surface of the front section  40  of the inner barrel  38 , is exposed and thus the total acoustic treatment area (provided by the inner surface of both the inner wall  50  and the front section  40  of the inner barrel  38 ) is increased, in contrast to the total acoustic treatment area substantially defined by only the inner surface of the inner wall  50  of the inlet cowl  44  when the inlet cowl  44  is in the retracted position. 
         [0026]    Optionally, one or more actuators  54  may be provided, for example being positioned in a space between the outer skin  36  and the inner barrel  38  of the main portion  32  and being supported on a stationary structure (not numbered) of the engine and may be operatively connected to the inlet cowl  44  for moving the inlet cowl  44  between the extended and retracted positions. 
         [0027]    Referring to FIGS.  1  and  4 - 7 , the nacelle  10  of the aeroengine may be configured to provide a variable geometry inlet system  11  b according to another embodiment. The description of the variable geometry inlet system  11  b below will be focussed on the structures and features which are different from those of the variable geometry inlet system  11  a described above and illustrated in  FIGS. 2-3 , and like structures and features will be indicated by like reference numerals and will not be redundantly described below. 
         [0028]    The fan assembly  14  may include a fan casing (not numbered) surrounding a fan rotor (not numbered). The fan casing according to this embodiment may be part of the inner barrel  38  of the main portion  32  of the nacelle  10  and may form a front section  40 ′ of the inner barrel  38 , positioned upstream of the front edge  42  of the annular outer skin  36  of the main portion  32 . The fan casing, at least a section thereof, thereby forms the stationary rear section of the inlet duct defined by the nacelle  10 . 
         [0029]    A plurality of circumferentially distributed plates  56  each may be pivotally connected about a tangential axis  58  (shown as a pivoting point in  FIGS. 4 and 5 ) at a front edge thereof, to an inside of the lip potion  46  of the inlet cowl  44 , to define an annular inner wall  50 ′. The circumferentially distributed plates  56  may overlap or interweave in the circumferential direction at adjacent side edges thereof. The interweaving or overlap of the plates  56  may be configured such that the annular inner wall  50 ′ defined by the plates  56  may be substantially free of gaps or ridges between adjacent plates  56 . An actuation system  60  may be provided to the plates  56  to actuate a pivotal motion about their respective tangential axis  58  such that the annular inner wall  50 ′ defined by the plates  56  is configured as a petal configuration having a truncated conical profile. The truncated conical profile may have a fixed diameter defined by the front edges of the plates  56  which corresponds to the inner diameter of the lip portion  46  of the inlet cowl  44 , and may have a variable conical angle resulting in a variable diameter at the rear edges (not numbered) of the plates opposite to the respective front edges of the respective plates. As the inlet cowl  44  is axially translated for example by the actuator  54  (only shown in  FIGS. 2 and 3 ), between the retracted and extended positions, the petal configuration of the annular inner wall  50 ′ may allow the conical angle thereof to be adjusted such that the annular front edge of the front section  40 ′ (the front edge of the fan casing in this embodiment) of the inner barrel  38  of the main portion  32 , is in contact with the respective plates  56 , thereby preventing formation of an abrupt step normal to the airflow  30 , between the plates  56  and the annular front section  40 ′ of the inner barrel  38  of the main portion  32 . This petal configuration also provides such a length-variable inlet duct by changing the conical angle thereof to allow a variable axial portion of the front section  40 ′ of the inner barrel  38  of the main portion  32 , to be inserted into the truncated conical profile of the plates  56 . 
         [0030]    When the inlet cowl  44  is in the retracted position and the front section  40 ′ of the inner barrel  38  is inserted deepest into the conical profile of the plates  56 , an axial portion of the plates  56  may be received in the annular space defined between the front section  40 ′ of the inner barrel  38  (the van casing in this embodiment) and the outer skin  48  of the inlet cowl  44 , as shown in  FIG. 4 . 
         [0031]    The plates  56  may be supported by a track system (which may be similar to those used for aircraft flaps) to ensure that the plates maintain their correct radial position throughout the translation of the plates  56  when the inlet cowl  44  moves between the retracted and extended positions. The actuation system  60  may include a series of individual actuators or a reduced number of actuators connected by a unison linkage system (not shown). The actuation system  60  may be made with electric, hydraulic or pneumatic means including an air motor or jack system which may be supplied by a branch from an inlet lip anti-icing system (not shown). 
         [0032]    The plates  56  may be provided with perforations therethrough or may have noise absorbing material applied on the inner surface thereof to thereby provide a variable-geometry acoustic treatment area which may increase when the inlet cowl  44  moves to the extended position to expose more inner surface of the plates  50 ′ (which form a portion of the axial length of the front section of the inlet duct) to the noise propagation through the inlet duct. 
         [0033]    It should be noted that the above embodiments of the described subject matter may be used to increase the inlet length/diameter ratio and thus the acoustic treatment area under aircraft landing approach conditions in order to achieve reduction of community noise. Nevertheless, the above-described subject matter may also be applicable to make a low drag inlet system for high bypass engines. In order to minimize drag with clean inlet conditions at high speed flight operation, the inlet and forward cowl should be short. However, during low speed flight operation, for example in landing approach, conditions such as cross winds and ground vortices can result in fan inlet flow distortion in the inlet. The variable-geometry inlet may be able to provide a high length/diameter ratio for those conditions and a short inlet with low forward cowl drag in high speed flight operation. 
         [0034]    A secondary outer skin  49  as shown in  FIGS. 2 and 3  may be optionally provided. The secondary skin  49  is a downstream extension of the annular outer skin  48  of the inlet cowl  44 , being stored within the nacelle  10  under the outer skin  36  of the main portion  32  when in the retracted position, and extending to cover a gap between the annular outer skin  48  and the and or outer skin  36  when in the extended position, in order to reduce external noise resulting from turbulence due to the discontinuity in the outer nacelle skin. The optional secondary outer skin  49  may be employed in situations where drag or turbulence need to be minimized for either noise or performance concerns. 
         [0035]    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 disclosed subject matter. For example, a turbofan aeroengine is described as an application of the described embodiments, however aeroengines of other types, such as pusher turboprop aeroengines or other may also be suitable for application of the described subject matter. Any suitable number of inlet sections may be provided. Any suitable relative motion, or combination of motions, may be used to apply the teachings hereof. 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.