Abstract:
An aircraft propulsion system has an engine assembly having an axis of rotation, a fan assembly operatively connected to the engine assembly and comprising a plurality of fan blades arranged circumferentially around the axis of rotation and a monolithic nacelle assembly which circumferentially encloses the fan assembly. The propulsion system has a ratio of fan assembly outer diameter to nacelle assembly outer diameter of at least 0.87. As a result, the propulsion system provides for a significantly larger fan assembly outer diameter for a given nacelle assembly outer diameter than a conventional prior art turbofan engine, thereby increasing the specific power output or the bypass ratio of the propulsion system, which will result in improved propulsive efficiency without an increase in nacelle drag.

Description:
This invention claims the benefit of UK Patent Application No. 1202790.0, filed on 20 Feb. 2012, which is hereby incorporated herein in its entirety. 
     FIELD OF THE INVENTION 
     This invention relates to aircraft propulsion systems and particularly, but not exclusively, to aircraft propulsion systems which utilize ducted turbofan type engines generally characterized by a high bypass ratio. 
     BACKGROUND TO THE INVENTION 
     As shown in  FIG. 1 , an axial flow gas turbine turbofan engine  10  comprises an air intake  11 , a low pressure compressor (or fan)  12 , an intermediate pressure compressor  13 , a high pressure compressor  14 , a combustor  15 , a high pressure turbine  16 , an intermediate pressure turbine  17 , a low pressure turbine  18 , and an exhaust nozzle  19 . A nacelle assembly  20  encloses the fan  12 . 
     In operation, air is drawn into the engine  10  through the intake  11  and accelerated by the fan  12 , to produce two air flows: a first air flow which enters the intermediate pressure compressor  13  and a second air flow which bypasses the core of the engine to provide direct propulsive thrust. 
     The ratio between the mass flow rates of these first and second air flows is termed the bypass ratio. 
     The first air flow entering the intermediate pressure compressor  13  is compressed before entering the high pressure compressor  14  where further compression takes place. 
     The compressed air leaving the high pressure compressor  14  is directed into the combustor  15  where it is mixed with fuel and the resulting mixture is combusted. The high pressure combustion products then rapidly expand as they pass through and drive the high, intermediate and low pressure turbines  16 ,  17  and  18 . The gas leaving the low pressure turbine  18  is then exhausted through the exhaust nozzle  19  and provides additional propulsive thrust. 
     The high, intermediate and low pressure turbines  16 ,  17  and  18  respectively drive the high and intermediate pressure compressors  14  and  13  and the fan  12  by means of separate interconnecting shafts. 
     It is known that increasing the bypass ratio of a 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. 
     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. 
     Most turbofan engines are designed to be capable of being mounted in an under-wing configuration. Such a configuration provides an upper limit to the nacelle diameter so as to maintain safe working ground clearance beneath the engine when it is installed on the aircraft. 
     Furthermore, turbofan engines typically have a relatively deep nacelle which encloses the fan and the core engine. This nacelle depth, in combination with the need to maintain a minimum safe ground clearance, limits the diameter of fan which can be employed on the engine. 
     STATEMENTS OF INVENTION 
     According to an aspect of the present invention there is provided an aircraft propulsion system nacelle comprising:
         an engine assembly; and   a fan assembly operatively connected to the engine assembly, and comprising a plurality of fan blades arranged circumferentially around an axis of rotation; and   a monolithic nacelle structure which circumferentially encloses the fan assembly, and comprises a first, radially proximal, surface locatable radially outward of the plurality of blades, and an opposite, second, radially distal, surface which forms a radially outermost surface of the nacelle assembly.       

     In the following description, the terms ‘forward’ and ‘rearward’ are to be understood to relate to air inlet and outlet portions of the nacelle assembly respectively, along the axis of rotation of the fan. In addition, the term ‘axial’ is understood to relate to the direction of the axis of rotation of the fan assembly. 
     A conventional turbofan engine nacelle employs a monocoque type construction and has a maximum nacelle depth of approximately 300 mm (12 in). The space within the nacelle is used to accommodate various engine ancillaries and a thrust reverser mechanism. Relocating these engine ancillaries to the core engine and providing reverse thrust provision through a variable pitch fan assembly enables the nacelle structure to be formed as a monolithic structure which can be made significantly thinner in the radial direction, and therefore lighter, than a conventional nacelle assembly. For example, a monolithic nacelle structure according to the present invention has a depth of only approximately 100 mm (4 in). 
     This means that for a given overall engine nacelle diameter, the engine of the present invention can employ a fan having a diameter that is approximately 400 mm (16 in) larger than that of a conventional turbofan engine. This increase in fan diameter results in the engine of the present invention having a significantly larger fan swept area, which can reduce fuel consumption, increase propulsive efficiency and reduce exhaust emissions. 
     The reduction in nacelle weight of the turbofan engine of the present invention means that the engine may have a larger diameter fan without any significant increase in overall engine weight or drag. This in turn means that no additional strengthening of aircraft engine mounting points is necessary to accommodate the engine of the present invention. Any additional drag would otherwise adversely affect aircraft efficiency (for example, through nett engine thrust acting on the engine pylon). 
     Optionally, the nacelle structure further comprises an integral, energy absorbing, containment portion positioned radially distal to the first surface of the nacelle structure. 
     To satisfy regulatory requirements, turbofan engines are required to demonstrate that if part or all of a fan blade were to become detached from the remainder of the fan, that the detached parts are suitably captured within an energy absorbing, containment system. 
     In a conventional turbofan engine, the containment system is removably secured to a radially inner facing surface of the fan casing by front and rear fasteners. 
     The use of an integral, energy absorbing, containment portion eliminates the need for fasteners, making the nacelle structure simpler, more compact and lighter. 
     Optionally, the nacelle structure is formed as a metal forging, with the metal being selected from the group comprising steel, titanium, aluminum and alloys thereof. 
     By forming the nacelle structure from a metal forging, its radial thickness may be smaller than a conventional nacelle structure formed with a monocoque construction. This allows for a correspondingly larger diameter fan assembly. 
     Optionally, the nacelle structure is formed from a fiber reinforced composite material, with the composite fiber being selected from the group comprising glass, carbon, boron, aramid and combinations thereof. 
     An advantage of using a fiber reinforced composite material to form the nacelle structure is that its weight may be reduced over a nacelle structure formed from a metallic material. However, since such composite materials have a lower density than metals, it is likely that the radial depth of a composite nacelle will be greater than that of a metallic nacelle. 
     Optionally, a radial thickness of the nacelle structure is greater at a first, inlet portion than at a plane of the fan assembly normal to the axis of rotation. 
     For a typical conventional turbofan nacelle structure, the radial thickness of the nacelle structure is greatest over the portion of axial length corresponding approximately to the fan blades. This ensures that the nacelle structure has sufficient internal volume to accommodate the various engine ancillaries and thrust reversing mechanisms which are normally located within the nacelle, whilst also providing sufficient mechanical strength and rigidity to contain a released fan blade. 
     A monolithic nacelle structure according to the invention can be made significantly thinner in the radial direction than a conventional nacelle structure, by virtue of the sandwich construction of an outer nacelle barrel and liner system within that provides structural stiffness and a containment function. 
     Optionally, the nacelle structure comprises one or more voids. 
     By forming a void, or cavity, on a radially distal surface of the nacelle structure, it is possible to accommodate one or more engine ancillaries, such as an engine control unit. 
     The presence of such a void may also reduce the weight of the nacelle structure, particularly in regions of the structure, such as the forward, inlet portion, which are of increased radial thickness. 
     Optionally, at least one of the one or more voids is filled with a material having a lower density than the material forming the nacelle structure. 
     The density of the void filler material may be dictated by the requirement to provide resistance to foreign object damage (FOD) and general structural duty requirements. 
     In one arrangement of the invention, any such voids are filled with a low density material, such as an expanded closed cell foam material, in order to minimize the additional weight of such void filler material. 
     Optionally, the fan assembly has a fan pressure ratio of less than 1.3. 
     In this context, the term ‘fan pressure ratio’ is defined as follows: 
     
       
         
           
             
               
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     A conventional turbofan aircraft engine typically has a fan pressure ratio that is greater than 1.5. 
     A benefit of a fan which has a low fan pressure ratio is that the fan blade tip speed is reduced, which in turn reduces the level of noise generated by the engine fan. 
     Optionally, a ratio of an outer diameter of the fan assembly to an outer diameter of the nacelle assembly is greater than 0.87. 
     A typical conventional turbofan aircraft engine with a 3.56 m (140 in) diameter fan has a nacelle outer diameter of approximately 4.47 m (176 in). Such an engine will have a ratio of an outer diameter of the fan assembly to an outer diameter of the nacelle assembly of approximately 0.79. 
     By decreasing the radial thickness of the nacelle structure it is possible to increase the fan diameter for a given nacelle structure outer diameter. 
     This may result in an engine having a higher thrust output (by virtue of the larger fan diameter) than a conventional turbofan engine of comparable nacelle structure outer diameter. 
     Alternatively, this may result in an engine having the same thrust output as a corresponding conventional turbofan engine, but with a lower pressure ratio, lower blade tip speed and higher bypass ratio (by virtue of the larger fan diameter). This in turn results in decreased noise level and increased engine efficiency without any increase in nacelle drag. 
     Optionally, the containment liner assembly comprises a plurality of energy absorbing panels, each panel extending axially along the first surface of the nacelle structure. 
     By arranging the containment system as a plurality of energy absorbing panels, it becomes easier and more convenient to install the liners into the nacelle structure to complete the containment system. 
     Optionally, at least one of the plurality of energy absorbing panels is individually replaceable. 
     In the event that the containment system is damaged by impact, it will be necessary to repair or replace only those individual panels which have been damaged. 
     In one embodiment of the invention, by removing one or more individual panels it becomes possible to remove a single, variable pitch fan blade without the need to dismantle the nacelle structure itself. This makes minor repair and servicing tasks which involve only a single fan blade more convenient and cost effective than prior art techniques. 
     Optionally, the aircraft propulsion system nacelle comprises a nacelle structure support, the nacelle structure support being attached to the engine assembly, and the nacelle structure being removably attached to the nacelle structure support to thereby expose the fan rotor assembly. 
     In an embodiment of the invention, the aircraft propulsion system nacelle is split laterally into a forward, nacelle structure and a rearward, nacelle structure support, with the plane of the split being normal to the axis of rotation. The rearward, nacelle structure support is removably attached to the forward, nacelle structure in order to provide access to the core engine for maintenance and repair purposes. 
     Optionally, the nacelle structure comprises an inlet diffuser. 
     In one embodiment of the invention, the forward, nacelle structure is formed with a flow controlled inlet diffuser. 
     The inlet diffuser decreases the velocity of the air flow entering the fan assembly whilst at the same time increasing the pressure of the air flow. The geometry of the inlet diffuser is important to smooth the air flow into the fan assembly in order to minimize flutter and forced vibration of the fan blades. 
     Optionally, a ratio of the combined length of the nacelle structure and nacelle structure support to an outer diameter of the fan assembly is less than 1.25. 
     In a conventional turbofan engine, the ratio of the length of the nacelle structure to an outer diameter of the fan assembly is typically greater than approximately 1.35 to 1.40. 
     In an embodiment of the present invention, by forming the nacelle structure as a monolithic component having a smaller radial thickness than a conventional nacelle structure, and particularly by eliminating the nacelle mounted thrust reversing assembly, it becomes possible to reduce the axial length of the nacelle structure. 
     The reduced length of the nacelle structure reduces the wetted area of the nacelle and thereby reduces the skin drag, thereby increasing the aerodynamic efficiency of the nacelle structure. 
     Optionally, the aircraft propulsion system further comprises a thrust reverser means operable to produce reverse thrust by selectively varying a pitch angle of the plurality of fan blades. 
     In a conventional, aircraft turbofan engine the thrust reversing facility is typically provided by a mechanism which is housed in a rearward portion of the nacelle structure. In order to reduce the radial thickness of the nacelle structure it becomes necessary to remove the thrust reversing mechanism from the nacelle. 
     In the present invention the use of a variable pitch fan (in which the fan blade pitch angle may be reversed) enables the fan assembly to direct the thrust forwards. 
     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 
       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: 
         FIG. 1  shows a schematic partial sectional view of a gas turbine engine according to the prior art; 
         FIG. 2  shows a schematic, partial sectional view of a high bypass ratio aircraft gas turbine engine according to the prior art; 
         FIG. 3  shows a schematic, partial sectional view of a high bypass ratio aircraft gas turbine engine according to an embodiment of the present invention; 
         FIG. 4  shows a schematic, partial sectional view of the nacelle of the engine of  FIG. 3 ; 
         FIG. 5  shows a schematic partial sectional view of the nacelle of  FIG. 4 ; 
         FIG. 6  shows a schematic, partial sectional radial view of the nacelle of  FIG. 4 ; and 
         FIGS. 7A and 7B  show schematic, partial sectional circumferential views of the nacelle of  FIG. 4  showing how the removable liner panel allows the removal of a fan blade from its hub bearing assembly (for a variable pitch fan blade). 
     
    
    
     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 
     Referring to  FIGS. 3 and 4 , an aircraft propulsion system nacelle according to a first embodiment of the invention is designated generally by the reference numeral  100 . 
     The aircraft propulsion system nacelle  100  comprises an engine assembly  110 , a fan assembly  120  which is operatively connected to the engine assembly  110 , and a nacelle structure  130  which circumferentially encloses the fan assembly  120 . The engine assembly  110  has an inlet  112  and an exhaust  114 . 
     In this embodiment of the invention, the engine assembly  110  is a gas turbine engine having a conventional three-shaft configuration and having an axis of rotation  124 . 
     In the following description, the term ‘axially’ is to be understood to relate to the direction of the axis of rotation  124 . Similarly, the terms ‘forward’ and ‘rearward’ are to be understood to refer to the inlet  112  and exhaust  114  ends of the engine assembly  110  respectively. 
     The fan assembly  120  comprises a plurality of fan blades  122  which are arranged circumferentially around the axis of rotation  124 . 
     The nacelle structure  130  is formed as a monolithic component and comprises a first, radially proximal surface  132  and a second, radially distal surface  134 . The first surface  132  is positioned immediately radially outward of an outer circumference  126  of the fan assembly  120 . The second surface  134  forms a radially outward surface of the nacelle structure  130 . 
     In the present embodiment the nacelle structure  130  is formed as a fiber reinforced composite barrel  156  having a first radially proximal surface  132  and a second radially distal surface  134 . 
     The nacelle structure  130  includes an energy absorbing, containment portion  160  which is located between the outer circumference  126  of the fan assembly  120  and the first surface  132  of the nacelle structure  130 . 
     In the present embodiment, the containment portion  160  is arranged to extend circumferentially around the axis of rotation  124 , and axially forward and rearward of the fan assembly  120 . A first, radially proximal surface  162  of the containment portion  160  is contiguous with the first, radially proximal surface  132  of the nacelle structure  130 , forward and rearward of the containment portion  160 . 
     As shown in  FIG. 5 , the containment portion  160  includes a layered radial arrangement of an abradable liner  167 , a layer of high density honeycomb material  166  and a layer of low density honeycomb material  165 . These layers  165 ,  166 ,  167  are bonded together (separated by carbon or glass fiber laminate) with the composite barrel  156  to form a sandwich structure  130 . 
     The abradable liner  167  is adjacent to the outer diameter of the fan assembly and provides a sacrificial surface against which the fan blades  122  may rub during normal operation to form a gas-tight seal. 
     The arrangement of low density and high density honeycomb materials provides a structurally stiff and compact sandwich structure. This sandwich structure is capable of absorbing the impact energy associated with the impact of ice or other foreign objects, or released fan blades  122 . 
     In the present embodiment, a portion of the abradable liner and the low and high density honeycomb materials  167 ,  166 ,  165  is formed as a removable panel  180 , as shown in  FIG. 6 , the panel  180  extending forward and rearward of the fan blades  122 . 
     The panel  180  is secured in place in the nacelle structure  130  by a plurality of fasteners  184  each of which threadingly engage with corresponding threaded inserts  182  located in the containment portion  160 . 
     The interface  188  between the panel  180  and the containment portion  160  is reinforced with additional layers of composite material to allow for repeated removal and replacement without adversely affecting the alignment of the panel  180  when installed in the containment portion  160 . 
     As shown in  FIGS. 7A and 7B , the panel  180  may be removed to facilitate the removal of a single fan blade  122 . The panel  180  is configured such that its circumferential length  190  is less than the pitch  123  between adjacent blades  122 , when the fan blade  122  is in a feathered configuration (see  FIG. 7A ). 
     The fan assembly  120  is positioned such that two adjacent blades  122 , in the feathered position, straddle the panel  180 . This enables the panel  180  to be detached and withdrawn from between the two fan blades  122 . 
     The fan assembly  120  is then rotated by half the fan blade pitch  123  to leave one of the fan blades  122  aligned with the space  181  left by the removal of the panel  180 . The fan blade  122  is then released from the fan hub  125  and moved radially outwards from the hub  125  into the space  181 . 
     A replacement fan blade  122  may then be installed in the hub  125 , the fan assembly rotated by half the fan blade pitch  123 , and the panel  180  replaced. 
     The nacelle structure  130  includes a recess or void  170 . The void  170  is filled with a void filler material  172 , such as, in this case, a honeycomb material or syntactic foam (such as Rohacell™). The void filler material  172  is then covered with an acoustic panel  174 , in this case an acoustic honeycomb with a perforate skin. This arrangement of filler material and acoustic panel  172 ,  174  provides structural reinforcement, FOD resistance and acts as an acoustic liner. 
     A radial thickness  136  of the nacelle structure  130  varies along an axial length  135  of the nacelle structure  130  such that the variation defines an aerofoil profile. 
     A forward portion of the first, radially proximal surface  132  of the nacelle structure  130  is formed as an inlet diffuser  144 . 
     The aircraft propulsion system nacelle  100  further comprises a nacelle structure support  142 , the nacelle structure  130  and the nacelle structure support  142 , being separated by a nacelle joint  146 . The nacelle joint  146  is oriented normally to the axis of rotation  124 . The nacelle structure  130  and nacelle structure support  142  each extend from the nacelle joint towards the inlet  112  and exhaust  114  of the engine assembly  110  respectively. 
     The nacelle structure support  142  is attached to the engine assembly  110 , while the nacelle structure  130  is removably attached to the nacelle structure support  142 . The nacelle structure  130  can be removed to provide maintenance and repair access to the fan assembly  120 . 
     In a conventional aircraft turbofan engine, as shown in  FIG. 2 , the nacelle assembly  20  is formed as a hollow monocoque structure and is used to contain several engine sub-systems, such as, for example, oil tank, heat exchangers, LP gearbox and thrust reverser mechanism. 
     In the embodiment of the present invention, these sub-systems have been eliminated or relocated to the internal core of the main engine assembly  110 . For example, the fuel/oil heat exchanger has been moved to the engine core, while the incorporation of a variable pitch mechanism into the fan assembly  120  enables the pitch of the fan blades to be reversed thereby providing a reverse thrust facility. This enables the radial thickness  136  of the nacelle structure  130  to be significantly reduced. 
     In the present embodiment, a ratio of an outer diameter  128  of the fan assembly  120  to an outer diameter  138  of the nacelle structure  130  is 0.90. In this arrangement, for a given nacelle outer diameter  138  the fan assembly diameter  128  is approximately 400 mm (16 in) larger than that of a conventional prior art turbofan engine  10 . 
     In addition, in the present embodiment, since the nacelle structure  130  is no longer required to contain a thrust reverser mechanism, the nacelle structure  130  may be made axially shorter than that of a conventional prior art turbofan engine  10 . 
     The nacelle structure  130  of the present embodiment has a ratio of nacelle assembly length  135  to fan assembly outer diameter  128  of 1.15. In contrast, the ratio of nacelle assembly length to fan assembly outer diameter of a conventional prior art turbofan engine is typically greater than 1.30. 
     The present invention may be embodied in other specific forms without departing from its 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.