Abstract:
An assembly useful in reducing aircraft engine noise such as turbofan engine noise comprises: (a) a core portion comprising at least one entrance end and at least one exit end for the passage of acoustic energy therethrough; (b) a first member having an exterior face and an interior face, wherein the first member is adjacent to at least a part of the core portion, and the first member has a plurality of openings therein to permit the passage of acoustic energy therethrough; and (c) a second member having an exterior face and an interior face, wherein the interior face of the second member is adjacent to the core portion. The core portion may be a honeycomb acoustic structure, and the first member may be perforated to permit the passage of acoustic energy into and out of the conduit portion.

Description:
BACKGROUND OF THE INVENTION 
   1. Field of the Invention 
   This invention is directed to an assembly, method and system for reducing aircraft noise generated by aircraft engines such as turbofan engines, and to a method of making such an assembly. More particularly, this invention relates to an assembly located in a portion of an aircraft such as a turbofan engine nacelle inlet, engine fan case, core cowl, thrust reverser, core casing, and/or center body. The assembly comprises a core portion having at least one entrance end and at least one exit end for the passage of acoustic energy therethrough. The assembly also comprises a first member having an exterior face and an interior face, wherein the first member is adjacent to at least a portion of the core, and the first member has a plurality of openings therein to permit the passage of acoustic energy therethrough. 
   2. Background Information 
   The desirability of reducing the noise generated by aircraft engines such as turbofan engines is well known to those skilled in the art. As disclosed, for example, in U.S. Pat. No. 6,112,514, one method of reducing such noise which has been proposed is the use of Herschel-Quincke (H-Q) tubes of appropriate length arranged in an array about a turbofan engine to reduce the noise levels generated by the engine. Such an array of tubes, if properly located about the engine, creates destructive energy waves that cancel the acoustic energy in the turbofan engine yet do not contribute to any significant aircraft drag or reduced fuel consumption. As is also disclosed in U.S. Pat. No. 6,112,514, various parameters such as tube length, cross-sectional area etc. may be controlled via a control system responsive to varying engine operational or environmental conditions. In such a control system, sound is reintroduced in an out-of-phase relation from the sound propagating from the engine fan to effect sound cancellation. Such a control system may employ feedback or feedforward control, or a combination thereof. WO 02/059474 discloses an assembly and method for reducing such noise using at least one dynamically adaptable H-Q tube which is capable of being dynamically adapted with respect to tube geometry and acoustical characteristics to optimize cancellation of the predominant source tone for different engine cycles. 
   Existing H-Q tube configurations such as those described in U.S. Pat. No. 6,112,514 and WO 02/059474 are attached to the external surface of aircraft components, and thus are limited to specific shapes which are not always amenable to integration with the structure of the engine. However, it would be useful to employ a noise reduction assembly, which is integrated into the structure of aircraft components such as the turbofan engine nacelle core cowl. Accordingly, the assemblies of this invention may be employed in areas of little or no clearance, such as the core cowl. In contrast, a prior art external H-Q tube cannot be employed within the core cowl, because the external H-Q tube will physically interfere with other components. 
   In addition, although the use of an array of H-Q tubes within a “passive liner treatment” has previously been disclosed, for example, in U.S. Pat. No. 6,112,514, such a configuration employs an array of tubes which is not substantially contained within the passive liner treatment. In contrast, according to the present invention, the acoustic energy passes through, for example, conduit or conduits which are substantially contained within the core portion, thereby achieving various structural and cost benefits. In addition, U.S. Pat. No. 6,112,514 does not disclose the use of a first member having a plurality of openings therein (such as a perforated skin) which is operatively employed with entrance and exit ends of the core portion to permit the passage of acoustic energy into and out of the conduit. Moreover, in U.S. Pat. No. 6,112,514 a screen is required at the ends of the tubes. In contrast, the assembly of this invention does not employ a screen at the entrance and exit ends of the core portion. 
   Accordingly, it is one object of this invention to provide a noise reduction assembly which may be integrated within the structure of aircraft components. It is another object of this invention to provide a method of reducing aircraft noise using such an assembly. It is yet another object of this invention to provide an aircraft noise reduction system which employs such an assembly. Other objects, features and advantages of the present invention will be apparent to those skilled in the art from the detailed description of the invention and its various embodiments as described herein. 
   SUMMARY OF THE INVENTION 
   According to the present invention, an assembly useful in reducing aircraft engine noise such as turbofan engine noise comprises:
         (a) a core portion comprising at least one entrance end and at least one exit end for the passage of acoustic energy therethrough;   (b) a first member having an exterior face and an interior face, wherein the first member is adjacent to at least a part of the core portion, and the first member has a plurality of openings therein to permit the passage of acoustic energy therethrough; and   (c) a second member having an exterior face and an interior face, wherein the interior face of the second member is adjacent to the core portion.       

   A system for reducing aircraft engine noise comprises:
         (a) an apparatus for generating non-uniform noise energy about an aircraft structure; and   (b) an assembly for reducing aircraft noise comprising:
           (i) a core portion comprising at least one entrance end and at least one exit end for the passage of acoustic energy therethrough,   (ii) a first member having an exterior face and an interior face, wherein the first member is adjacent to at least a part of the core portion, and the first member has a plurality of openings therein to permit the passage of acoustic energy therethrough, and   (iii) a second member having an exterior face and an interior face, wherein the interior face of the second member is adjacent to the core portion.   
               

   In one embodiment, the core portion is a honeycomb acoustic structure having a conduit therein, and the first member is perforated to permit the passage of acoustic energy into and out of the conduit. 
   In another embodiment, a plurality of conduits are located within the assembly, the core portion is a honeycomb acoustic structure which substantially contains the conduits, and the first member is perforated to permit the passage of acoustic energy into and out of the conduits. 
   A method of making an assembly for reducing aircraft engine noise is also described. The method comprises:
         (a) providing a support member having a plurality of vertical walls;   (b) providing a plurality of enclosure members having a plurality of vertical walls;   (c) forming a subassembly by having at least one vertical wall of at least one enclosure member adjoin at least one vertical wall of at least one support member;   (d) providing a core portion having an aperture, a first face, and a second face;   (e) providing a first member having an exterior face and interior face, wherein the interior face of the first member is adjacent to at least a part of the first face of the core portion, and the first member has a plurality of openings therein to permit the passage of acoustic energy therethrough;   (f) providing the subassembly into the core aperture; and   (g) providing a second member adjacent to the second face of the core portion, such that the subassembly and second member define a conduit to permit the passage of acoustic energy therethrough.       

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  depicts a prior art arrangement of H-Q tubes in a turbofan aircraft engine. 
       FIGS. 2A–2D  depicts prior art arrangements of H-Q tubes in a turbofan aircraft engine. 
       FIG. 3  depicts a turbofan aircraft engine of the type in which the assembly of this invention may be employed. 
       FIG. 4A  depicts a cross-sectional view of one embodiment of the assembly of this invention. 
       FIG. 4B  depicts a cross-sectional view of another embodiment of the assembly of this invention. 
       FIG. 5  depicts a cross-sectional view of another embodiment of the assembly of this invention. 
       FIG. 6  depicts an exploded view of the embodiment depicted in  FIG. 4B . 
       FIG. 7  depicts a cross-sectional view of another embodiment of the assembly of this invention. 
       FIGS. 8A and 8B  depict cross-sectional and bottom views, respectively, of the conduit-defining member used in the assembly of this invention depicted in  FIGS. 4A and 4B . 
       FIGS. 9A and 9B  depict cross-sectional and bottom views, respectively, of another embodiment of the conduit-defining member which may be used in the assembly of this invention. 
       FIG. 10A  depicts an exploded view of an assembly of this invention using the conduit portion of  FIGS. 8A and 8B . 
       FIG. 10B  depicts an exploded view of an assembly of this invention using the conduit portion of  FIGS. 9A and 9B . 
       FIGS. 11A and 11B  depict a support member and enclosure members which are used to form one embodiment of the assembly of this invention. 
       FIGS. 12A and 12B  depict the formation of a subassembly using the support member and enclosure members of  FIGS. 11A and 11B . 
       FIGS. 13A and 13B  depict the formation of one embodiment of the assembly of this invention using the subassembly of  FIGS. 12A and 12B . 
       FIGS. 14A and 14B  depict cross sectional views of other embodiments of the assembly of this invention. 
   

   DETAILED DESCRIPTION OF THE INVENTION 
   This invention is directed to an assembly, method and system for reducing fan noise from a noise generating system such as a turbofan engine, and to a method of preparing such an assembly. In a preferred embodiment, the assembly, method and system of this invention are used with a turbofan engine. However, other systems such as, for example, systems that generate noise which propagates in a partially enclosed area, such as air ventilation ducts, are equally contemplated for use with the present invention. Accordingly, the assembly, method and apparatus of this invention are not limited to use in conjunction with turbofan engines only, but instead may be used with other noise generating devices or systems. The dimensions of the assembly of this invention, including acoustic energy path length, width, shape and other variables and quantities specified herein may vary depending on the particular application of this invention. Accordingly, numbers and dimensions specified herein are not to be construed as limitations on the scope of this invention, but are meant to be merely illustrative of one particular application. 
   For exemplary purposes only, the noise reduction system of the present invention is described with reference to reducing noise in turbofan engines. According to this exemplary embodiment, the noise reduction system of the present invention effectively reduces noise energy over a wide range of frequencies for both tonal and broadband components of the inlet and outlet noise for turbofan engines. Specifically, the noise reduction system of the present invention utilizes at least one assembly preferably comprising at least one conduit, more preferably a plurality of such conduits arranged in a series or parallel array substantially within an acoustic panel within a turbofan engine to reduce the noise levels generated by the turbofan engine. The assembly or assemblies may also be placed at the inlet, and in other locations, such as, for example, in either the upstream or downstream locations of the turbofan engine. The inlet and outlet of the conduits of the assembly can be placed parallel to the engine axis or at an angle. By using an array of such assemblies in the turbofan engine nacelle inlet (or other noise generating systems), destructive waves are created that cancel the acoustic energy in the turbofan engine, without contributing to any significant aircraft drag or reduced fuel consumption. 
     FIG. 1  and  FIGS. 2A–2D  are illustrative of the operation of prior art H-Q tube systems, as set forth and described in U.S. Pat. No. 6,112,514, incorporated herein by reference. In FIGS.  1  and  2 A– 2 D, the H-Q tubes are tubular passageways attached to the exterior surface of the annular duct structure and penetrate through the duct walls. 
     FIG. 3  depicts a cross-sectional view of a turbofan aircraft engine of the type in which the assembly of this invention may be employed. In  FIG. 3 , engine  300  has nacelle  304 , which includes nacelle inlet  324 , core section  326 , primary exhaust section  328 , fan cowl  368 , and core cowl  366 . Located downstream from the nacelle inlet  324  are fan  332  and stator  333 . Fan  332  sends bypass air through fan duct  334  which is exhausted via fan duct outlet  336 . The assembly of this invention may be employed, for example, in the nacelle inlet  324  (defined by the region between areas A and B), the engine fan case  364  (defined by the region between areas B and C), the core cowl  366  (defined by the region between areas F and G), the nacelle-thrust reverser  370  (defined by the region between areas C and D), the engine core casing  367  (defined by the region between areas E and F), and/or center body  360  (defined by the region between areas H and I). 
     FIG. 4A  depicts a cross-sectional view of one embodiment of this invention, in which the embodiment depicted is a cross-sectional view of the nacelle portion designated “J” in  FIG. 3 . In  FIG. 4A , the nacelle  304  contains a backskin or panel  406 , a core assembly  408 , and a perforated skin or panel  410 , which forms at least a part of the inner wall of nacelle inlet  324 , as shown. Core assembly  408  contains a cavity  401  which in turn contains a conduit-defining member  412 , having a conduit  403  therein. Conduit  403  is defined by outer wall  416  and inner wall  418  and has an entrance end  411  and an exit end  413 . As depicted in  FIG. 4A , inner wall  418  is essentially parallel to the inner wall (which in this embodiment is perforated skin  410 ) of nacelle inlet  324 , and outer wall  416  is essentially parallel to backskin  406 . Conduit-defining member  412  may be fabricated from aluminum, graphite or other suitable materials, and is typically fabricated from the same material as panels  406  or  410 . Core assembly  408  is typically a honeycomb structure and may be divided into an upper portion  414  and a lower portion  415  by a septum  405 . As is well known to those skilled in the art, the septum separates layers within a honeycomb core, thereby improving the acoustic performance of a honeycomb panel structure. The septum  405  may be fabricated from a metallic sheet having perforations, a metallic mesh material, or a nonmetallic mesh material such as a polymeric mesh material. The use of core assemblies and panel structures for aerospace applications are well known to those skilled in the art, as exemplified by the honeycomb-shaped core panel structures depicted in U.S. Pat. Nos. 4,869,421, 5,445,861 and EP Application No. 1238741. The core assembly may be made of a titanium alloy material such as titanium-aluminide, and the backskin and perforated skin may also be made of a titanium alloy such as a titanium-aluminide facing sheet. The core may have a honeycomb shape, or may have other geometric shapes. The core assembly, backskin and perforated skin may be joined via conventional bonding techniques such as liquid interface diffusion (LID) bonding, brazing and adhesive bonding. 
   As shown, for example, in  FIG. 4A , during operation of the engine, acoustic energy (indicated by arrows) passes through perforated skin  410 , enters entrance end  411 , and exits via exit end  413 . The acoustic energy behaves as a pressure wave, and the pressure differential causes acoustic energy to enter via entrance end  411  and exit via exit end  413 . The conduit defined by the entrance end  411  and exit end  413  therefore functions as an acoustic wave guide. In a preferred embodiment, during operation of the engine the flow of air through the conduit is  de minimis . In a particularly preferred embodiment, in which the assembly of this invention is located in the engine inlet or forward of the fan, no air passes through the conduit. However, if the assembly is located aft of the fan, such as in the by-pass duct, air may pass through the conduit. This is also true for the additional embodiments of the invention described herein. 
     FIG. 4B  depicts a cross-sectional view of another embodiment of this invention, in which the embodiment depicted is a cross-sectional view of the nacelle portion designated “J” in  FIG. 3 . In  FIG. 4B , nacelle  304  contains a backskin or panel  456 , a core assembly  458 , and a perforated skin or panel  460  which forms at least a part of the inner wall of nacelle inlet  324 , as shown. Core assembly  458  contains a cavity  451  which in turn contains a conduit-defining member  462 , having a conduit  453  therein. Conduit  453  is defined by outer wall  466  and inner wall  468  and has an entrance end  461  and an exit end  463 . As depicted in  FIG. 4B , inner wall  468  is essentially parallel to the inner wall (which in this embodiment is perforated skin  460 ) of nacelle inlet  324 , and outer wall  466  is essentially parallel to backskin  456 . Conduit-defining member  462  may be fabricated from materials as previously described with respect to conduit-defining member  412  in  FIG. 4A . In the embodiment depicted in  FIG. 4B , septum  455  divides core assembly  458  into an upper portion  464  and a lower portion  465 . The core assembly, backskin, perforated skin and septum may be fabricated from materials as previously described with respect to the embodiment depicted in  FIG. 4A . However, unlike the embodiment depicted in  FIG. 4A , in  FIG. 4B  the septum  455  extends continuously through the core assembly  458 , as shown. Accordingly, during operation of the engine, acoustic energy (indicated by arrows) passes through perforated skin  460  and septum  455 , enters entrance end  461  and exits through septum  455  and exit end  463 . 
   In preferred embodiments of this invention as depicted in  FIGS. 4A and 4B , substantial portions of the core reside between the conduit and the perforated skin or panel. For example, in  FIG. 4A , lower core portion  415  resides between conduit  403  and perforated skin or panel  410 , and in  FIG. 4B , lower core portion  465  resides between conduit  453  and perforated skin or panel  460 . Accordingly, the conduit  403  is integrated within the core. 
     FIG. 5  depicts a cross-sectional view of another embodiment of this invention, in which the embodiment depicted is a cross-sectional view of the nacelle portion designated “J” in  FIG. 3 . In  FIG. 5 , nacelle  304  contains a perforated skin or panel  510 , a backskin or panel  506  which forms at least a part of the inner wall of nacelle inlet  324 , as shown, and a honeycomb-shaped core  508  having an upper portion  514  and a lower portion  515  separated by a septum  505 . Core  508  contains a conduit-defining member  512  which in this embodiment is machined or otherwise formed or fabricated from a material impervious to acoustic energy such as aluminum or another suitable material and placed in a cavity formed in core  508 . Conduit-defining member  512  has a conduit  503  therein. The conduit  503  is defined by outer wall  516  and inner wall  518 , and has an entrance end  511  and exit end  513 . Acoustic energy (indicated by arrows) passes through perforated skin  510 , enters conduit  503  via entrance end  511 , and exits via exit end  513 . As depicted in  FIG. 5 , inner wall  518  is essentially parallel to the inner wall (which in this embodiment is perforated skin  510 ) of nacelle inlet  324 , and outer wall  516  is substantially parallel to backskin  506 . In another embodiment, inner wall  518  may be acoustically treated, such as by adding perforations to improve system acoustic performance. 
     FIG. 6  depicts an exploded view of the embodiment of  FIG. 4B . As shown in  FIG. 6 , the assembly is made up of a backskin  406 , an upper core portion  414  having a cavity  401  therein, a septum  405 , a lower core portion  415 , and a perforated skin  410 . A conduit-defining member  412  having an entrance end  413  and exit end  411  resides within cavity  401 . In another embodiment (not shown), the septum may be locally removed in the region of cavity  401 . 
     FIG. 7  depicts another embodiment of this invention, in which the embodiment depicted is a cross-sectional view of the nacelle portion designated “J” in  FIG. 3 . In  FIG. 7 , the assembly is made up of a backskin  706 , a honeycomb-shaped core  708 , and a perforated skin  710  which forms at least a part of the inner wall of nacelle inlet  324 , as shown. A septum (not shown in  FIG. 7 ) may optionally be employed. In this embodiment, conduit-defining member  712  is machined or otherwise formed or fabricated from a solid material such as aluminum or another suitable acoustically impervious material and placed in a cavity formed in core  708 . Conduit-defining member  712  contains a conduit  703  defined by outer wall  716  and inner wall  718 , and having an entrance end  711  and multiple exit ends  713  and  715 . As depicted in  FIG. 7 , inner wall  718  is essentially parallel to the inner wall (which in this embodiment is perforated skin  710 ) of nacelle inlet  324 , and outer wall  716  is substantially parallel to backskin  706 . Acoustic energy (indicated by arrows) passes through perforated skin  710 , enters conduit  703  via entrance end  711 , and exits via exit ends  713  and  715 . The presence of more than one exit end provides enhanced acoustic performance by permitting two tones or noise frequencies to be attenuated, thereby enhancing the noise reduction achieved using this invention. 
     FIGS. 8A and 8B  depict cross-sectional and bottom views, respectively, of the configuration of the conduit-defining member portion  412  depicted in  FIG. 4A . In  FIGS. 8A and 8B , conduit defining-member  412  is fabricated such that it has apertures  411  and  413  in floor portion  420 . Conduit  403  is defined by outer wall  416  and inner wall  418 . Apertures  413  and  411  act as entrance and exit ends, respectively, as previously described. 
     FIGS. 9A and 9B  depict cross-sectional and bottom views, respectively, of another embodiment of a conduit-defining member which may be used in this invention. Conduit-defining member  912  is fabricated such that it has apertures  911  and  913  in floor portion  922 . Apertures  913  and  911  act as entrance and exit ends, respectively, as previously described. In  FIG. 9A , the entire ceiling portion  920  is open, although in other embodiments only a part of the ceiling portion may be open. Conduit-defining member  912  may be placed in cavity  401  of core  408  (as previously described with respect to  FIG. 4A ), and a backskin or panel  406  may be placed adjacent to ceiling portion  920  of conduit-defining member  912 , such that backskin or panel  406  and floor portion  922  define conduit  903  in conduit-defining member  912 . 
     FIG. 10A  depicts an exploded view of an embodiment of the assembly of this invention using conduit-defining member portion  412 , as depicted in  FIG. 4A . As shown in  FIG. 10A , the assembly contains backskin  406 , upper core  414  having cavity  401  therein, conduit-defining member  412  which resides in cavity  401  (when assembled), lower core  415  and perforated skin  410 . A septum (not shown in  FIG. 10A ) may optionally be employed. 
     FIG. 10B  depicts an exploded view of an embodiment of the assembly of this invention using conduit-defining member  912 , as depicted in  FIGS. 9A and 9B . As shown in  FIG. 10B , the assembly contains backskin  406 , upper core  414  having cavity  401  therein, conduit-defining member  912  which resides in cavity  401  (when assembled), lower core  415  and perforated skin  410 . A septum (not shown in  FIG. 10B ) may optionally be employed. 
     FIGS. 11A ,  11 B,  12 A,  12 B,  13 A and  13 B set forth one embodiment of a method of making the assembly of this invention.  FIG. 11A  depicts a support member  1102  having a plurality of vertical walls  1104 . Support member  1102  and vertical walls  1104  may be fabricated from aluminum or another suitable rigid material.  FIG. 11B  depicts a plurality of enclosure members  1106  which may also be manufactured from aluminum or another suitable material. In this embodiment, the enclosure members  1106  are approximately elliptical in shape. Enclosure members  1106  may be fabricated from a single strip which is then bent or formed to the desired shape, where interface  1108  is the interface between the two ends of the strip. In other embodiments, enclosure members  1106  may be continuous pieces without interface  1108 . 
     FIGS. 12A and 12B  depict the preparation of the support member  1102  and enclosure members  1106  to form a subassembly  1110 . As shown, the support member  1102  and enclosure members  1106  are joined to each other to form the subassembly  1110 . The support member  1102  and enclosure members  1106  may be joined by conventional means, such as LID bonding. 
     FIG. 13A  depicts preparation of another embodiment of this invention employing subassembly  1110  as shown in  FIG. 12B . In  FIG. 13A , core portion  1112 , which may be a honeycomb material as previously described, is shown having a cavity  1114  therein. Septum  1116  (which optionally may be removed in the region of cavity  1114 ) is adjacent to core portion  1112 , and perforated skin  1118  is adjacent to septum  1116 . As shown in  FIG. 13B , subassembly  1110  is placed within the cavity or aperture  1114  such that subassembly  1110  is adjacent to septum  1116  (or perforated skin  1118  if the septum  1116  has been locally removed in the region of cavity  1114 ). Subassembly  1110  forms a plurality of conduit-defining members of the type depicted in  FIGS. 9A and 9B . A backskin (not shown in  FIGS. 13A  or  13 B) is thereafter applied to face  1120  of core  1112  to form the final assembly, which may be employed, for example, in an engine case liner or engine core cowl structure. This structure results in a plurality of adjacent conduit-defining members arranged adjacent to one another, such that when the structure is installed in a gas turbine engine, the conduit-defining members have the major axes thereof aligned parallel to the engine major axis. 
   The assembly of this invention may extend circumferentially about nacelle inlet  324  as depicted in  FIG. 3 . In other embodiments, the assembly may also be positioned in other locations, such as the bypass duct or primary nozzle. It will be understood by those skilled in the art that the assembly may extend or be axially located forward or aft of nacelle portion J shown in  FIG. 3 . For example, assemblies of this invention may be located in the core cowl  366 , the engine core casing  367 , and/or the center body  360 . 
   It will also be understood by those skilled in the art that more than one assembly as described herein may be employed within a given engine nacelle inlet, engine fan case, nacelle core cowl, engine core casing, nacelle center body, etc. 
   It will also be understood by those skilled in the art that the location of the exit end or ends relative to the entrance end in the various embodiments of the assembly of this invention may be asymmetrically varied longitudinally or axially, and that the entrance end and/or exit end or ends may be angled or offset with respect to the engine axis. For example,  FIG. 14B  depicts a core assembly  1458  having a conduit  1453  with entrance end  1461  and exit end  1463 , which are asymmetrically positioned with respect to center line CL, as shown. It will also be understood by those skilled in the art that the cross-sectional areas of the entrance and exit ends may be unequal. For example,  FIG. 14A  depicts a core assembly  1408  having a conduit  1403 , with entrance end  1411  and exit end  1413 . Entrance end  1411  and exit end  1413  have unequal cross-sectional areas. It will also be understood that the cross-sectioned shapes of the entrance and exit ends may be circular, elliptical or any other shape that allows acoustic energy to pass therethrough. 
   Without wishing to be bound by any one theory, it is expected that in this invention non-uniform noise energy is generated during operation of the engine. The assembly of this invention is employed to allow such non-uniform noise energy to pass therethrough, and to be reintroduced into the non-uniform noise energy in an out-of-phase relation to the non-uniform noise energy itself to effect sound cancellation and noise abatement. 
   While the invention has been described in terms of its preferred embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the appended claims.