Patent Publication Number: US-2023133478-A1

Title: Adaptive vertical lift engine (avle) fan

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application is a divisional application of, and claims priority under 35 USC § 120 to U.S. nonprovisional application Ser. No. 16/397,854, filed Apr. 29, 2019, the entire contents of which are incorporated by reference. 
    
    
     GOVERNMENT RIGHTS 
     This present application was made with United States government support under Contract No. W911W6-16-2-0011, awarded by the United States Army. The United States government has certain rights in the present application. 
    
    
     BACKGROUND 
     Fluid propulsion devices achieve thrust by imparting momentum to a fluid called the propellant. An air-breathing engine, as the name implies, uses the atmosphere for most of its propellant. The gas turbine produces high-temperature gas which may be used either to generate power for a propeller, fan, generator or other mechanical apparatus or to develop thrust directly by expansion and acceleration of the hot gas in a nozzle. In any case, an air breathing engine continuously draws air from the atmosphere, compresses it, adds energy in the form of heat, and then expands it in order to convert the added energy to shaft work or jet kinetic energy. Thus, in addition to acting as propellant, the air acts as the working fluid in a thermodynamic process in which a fraction of the energy is made available for propulsive purposes or work. 
     Typically, turbofan engines include at least two air streams. All air utilized by the engine initially passes through a fan, and then it is split into the two air streams. The inner air stream is referred to as core air and passes into the compressor portion of the engine, where it is compressed. This air is fed to the combustor portion of the engine where it is mixed with fuel and the fuel is combusted. The combustion gases are then expanded through the turbine portion of the engine, which extracts energy from the hot combustion gases, the extracted energy being used to run the compressor, the fan and other accessory systems. The remaining hot gases then flow into the exhaust portion of the engine, which may be used to produce thrust for forward motion to the aircraft. 
     The outer air flow stream bypasses the engine core and is pressurized by the fan. Typically, no other work is done on the outer air flow stream which continues axially down the engine but outside the core. The bypass air flow stream also can be used to accomplish aircraft cooling by the introduction of heat exchangers in the fan stream. Downstream of the turbine, the outer air flow stream is used to cool engine hardware in the exhaust system. When additional thrust is required (demanded), some of the fans bypass air flow stream may be redirected to the augmenter (afterburner) where it is mixed with core flow and fuel to provide the additional thrust to move the aircraft. 
     Many current and most future aircrafts need efficient installed propulsion system performance capabilities at diverse flight conditions and over widely varying power settings for a variety of missions. Current turbofan engines are limited in their capabilities to supply this type of mission adaptive performance, in great part due to the fundamental operating characteristics of their core systems which have limited flexibility in load shifting between shaft and fan loading. 
     When defining a conventional engine cycle and configuration for a mixed mission application, compromises have to be made in the selection of fan pressure ratio, bypass ratio, and overall pressure ratio to allow a reasonably sized engine to operate effectively. In particular, the fan pressure ratio and related bypass ratio selection needed to obtain a reasonably sized engine capable of developing the thrusts needed for combat maneuvers are non-optimum for efficient low speed flight where a significant portion of the engine output is transmitted to the shaft. Engine performance may suffer due to the bypass/core pressure leakage that may occur at reduced fan power/load settings. 
     Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,  FIG.  1   a    shows a general orientation of a turbofan engine in a cut away view. In the turbofan engine shown, the flow of the air is generally axial. The engine direction along the axis is generally defined using the terms “upstream” and “downstream” generally which refer to a position in a jet engine in relation to the ambient air inlet and the engine exhaust at the back of the engine. For example, the inlet fan is upstream of the combustion chamber. Likewise, the terms “fore” and “aft” generally refer to a position in relation to the ambient air inlet and the engine exhaust nozzle. Additionally, outward/outboard and inward/inboard refer to the radial direction. For example, the bypass duct is outboard the core duct. The ducts are generally circular and co-axial with each other. 
     As ambient inlet airflow  12  enters inlet fan duct  14  of turbofan engine  10 , through the guide vanes  15 , passes by fan spinner  16  and through fan rotor (fan blade)  42 . The airflow  12  is split into primary (core) flow stream  28  and bypass flow stream  30  by upstream splitter  24  and downstream splitter  25 . In  FIG.  2   , the bypass flow stream  30  along with the core/primary flow stream  28  is shown, the bypass stream  30  being outboard of the core stream  28 . The inward portion of the bypass steam  30  and the outward portion of the core streams are partially defined by the splitters upstream of the compressor  26 . The fan  42  has a plurality of fan blades. 
     As shown in  FIGS.  1   a  and  1   b    the fan blade  42  shown is rotating about the engine axis into the page, therefor the low pressure side of the blade  42  is shown, the high pressure side being on the opposite side. The Primary flow stream  28  flows through compressor  26  that compresses the air to a higher pressure. The compressed air typically passes through an outlet guide vane to straighten the airflow and eliminate swirling motion or turbulence, a diffuser where air velocity decreases, and a compressor manifold to distribute the air in a smooth flow. The core flow stream  28  is then mixed with fuel in combustion chamber  36  and the mixture is ignited and burned. The resultant combustion products flow through turbines  38  that extract energy from the combustion gases to turn fan rotor  42 , compressor  26  and any shaft work by way of turbine shaft  40 . While  FIGS.  1   a  and  1   b    only show one shaft for clarity, commonly turbine engines have multiple shafts or spools (e.g. high pressure spool, low pressure spool, etc.) The gases, passing exhaust cone, expand through an exhaust nozzle  43  to produce thrust. Primary flow stream  28  leaves the engine at a higher velocity than when it entered. Bypass flow stream  30  flows through fan rotor  42 , flows by bypass duct outer wall  27 , an annular duct concentric with the core engine flows through fan discharge outlet and is expanded through an exhaust nozzle to produce additional thrust. Turbofan engine  10  has a generally longitudinally extending centerline represented by engine axis  46 . 
     Current conventionally bladed core engines cannot maintain constant or near constant operating pressure ratios as bypass flow is reduced. Current conventionally bladed fan rotors do not have the flexibility in efficiently reducing fan pressure ratio while maintaining core pressure. 
     With reduced or no flow in the Bypass stream  30 , the core stream  28  relative pressure is greater than that in the Bypass stream  30 . In the area of the fan shown as  50  in  FIG.  1   b   , pressure differences between the core duct and the bypass duct can cause cross flow between the ducts in the area of the fan blade across the region  50  from the core stream  28  into the bypass stream  30  thus reducing the core pressure which has a deleterious effect on the operation of the core and un-necessarily loading the turbine to recover the lost pressure. 
     A partial blade splitter, similar to a partial span shroud or clapper, separating the core and bypass streams as described herein, can limit the pressure loss in the core and the subsequent degradation in output of the core engine while maintaining communication across the flows. The split flow path enables the fan to operate effectively in a turbofan mode and a turboshaft mode where the bypass flow, pressure and thrust are substantially reduced and power is available to the shaft. 
     These and many other advantages of the present subject matter will be readily apparent to one skilled in the art to which the invention pertains from a perusal of the claims, the appended drawings, and the following detailed description of preferred embodiments. 
     SUMMARY 
     According to some aspects of the disclosure, a turbofan engine may have a fan in fluid communication with a core stream and a bypass stream of air. The core stream may be compressed by the fan and a core compressor portion, heated and expanded through a core turbine portion. The core turbine portion may drive the fan and the compressor portion The core turbine portion may be connected to a shaft. The bypass stream may be compressed by the fan. The core and the bypass streams may be separated by a partial midspan shroud on the fan and a downstream splitter. The partial midspan shroud may extend axially forward from the trailing edge of the fan to at least the midchord of the fan. A seal between the trailing edge of the partial midspan shroud and the leading edge of the downstream splitter may restrict flow between the core stream and the bypass stream. 
     Some embodiments may include, an adjustable inlet guide vane upstream of the fan, the adjustable inlet guide vane positional between a first position and a second position, the second position restricting flow of the bypass stream more than the first position. It is envisioned, however not required, that the adjustable portion of the inlet guide vane only operates on the flow associated with the bypass stream and the portion of the guide vane in the flow associated with the core is fixed or independently adjustable. Some embodiments may include an upstream splitter on the adjustable guide vane, the upstream splitter having a trailing edge axially displaced from the leading edge of the fan. In some embodiments, the seal may be selected from the group consisting of labyrinth seal, lip seal and carbon seal. 
     In some embodiments, the midspan partial shroud extends axially forward from the trailing edge of the fan no more than ⅔ of a local chord on the fan. In some embodiments, the midspan partial shroud extends axially forward from the trailing edge of the fan no more than ⅞ to ½ of a local chord on the fan. The fan may have a blade span and the midspan partial shroud may be radially located on the middle third of the blade span. In some embodiments the midspan partial shroud may be concentric with the fan. 
     Some embodiments may include an additional splitter, a second fan, and a second seal, the second fan positioned upstream of the first fan, said second fan comprising a second midspan partial shroud extending axially at least to a local midchord of the second fan but short of the leading edge of the second fan, the second seal connecting the trailing edge of the second midspan partial shroud with the leading edge of the additional splitter. In some embodiments, the partial midspan shroud rotates about the engine axis with respect to the downstream splitter. 
     According to some aspects of the disclosure, a turbofan engine may include a fan, a core duct which may define a portion of a core fluid path, a bypass duct which may define a portion of a bypass fluid path. The bypass duct may be concentric with the core duct and radially displaced from the core duct. A downstream splitter may define an annular border portion between the core duct and the bypass duct, and downstream of the fan. An annular border region may extend between a leading edge of the fan and a leading edge of the downstream splitter. The annular border region may separate the core fluid path and the bypass fluid path. The fan may rotate through the annular border region. A shroud within the annular border region may extend between blades in the fan. The shroud may have a leading edge downstream from the leading edge of the blades and upstream of the midchord. The shroud may rotate with respect to the downstream splitter. A seal between a trailing edge of the shroud and the leading edge of the downstream splitter may restrict migration from the core fluid path to the bypass fluid path. A variable inlet guide vane upstream of the fan may restrict the bypass flow at a first position and may not restrict the bypass flow at a second position. 
     In some embodiments, a pressure in the core fluid path may be higher than a second pressure in the bypass fluid path when the variable inlet guide vane is at the first position. Some embodiments may include an upstream splitter defining an annular first border portion between the core duct and the bypass duct. In some embodiments, the seal is selected from the group consisting of labyrinth seal, lip seal and carbon seal. In some embodiments the shroud extends axially forward from the trailing edge of the fan no more than ⅔ of a local chord on the fan. In some embodiments, the shroud extends axially forward from the trailing edge of the fan no more than ⅞ to ½ of a local chord on the fan. In some embodiments, the fan has a blade span and the shroud may be radially located on the middle third of the blade span. In some embodiments, the shroud may be concentric with the fan. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIGS.  1   a  and  1   b    are illustrations representing conventional turbofan engines. 
         FIG.  2    is an illustration of the Bypass and primary stream flow paths. 
         FIG.  3    is an illustration of a turbofan engine according to an embodiment of the disclosed subject matter. 
         FIG.  4    is an illustration of a turbofan engine without an upstream splitter according to an embodiment of the disclosed subject matter. 
         FIG.  5    is an illustration of with multiple fan stages according to embodiments of the disclosed subject matter. 
         FIG.  6    is an illustration of the blade splitter according to embodiments of the disclosure subject matter. 
     
    
    
     DETAILED DESCRIPTION 
       FIG.  3    illustrates a Bypass flow duct  31  lying radially outward from the core flow duct  29 . The fan  42  is positioned upstream from the splitter  25  that separates air flow between the ducts. The inlet guide vane splitter  24  is positioned upstream from the fan  42  at radially inward of the adjustable inlet guide vane  15 . As the inlet guide vane  15  angle is changed, the bypass flow may be inhibited and pressure within the bypass flow duct  31  may differ from the pressure present in the core flow duct  29 . In prior systems, air can cross between the two ducts in the vicinity of the fan blade in region  50  as shown in  FIG.  1   b    thus causing detrimental engine performance in the core as described previously. 
       FIG.  3    illustrates a blade splitter  26  within the fan  42  and a splitter  25  behind the fan  42 . The splitter assembly  27  (blade splitter  26  and splitter  25 ) interface with each other with a rotating seal or discourager located just behind the fan  42 . The blade splitter  26  extends axially at least past the midchord  43  of the fan  42 . It is advantageous to have a long enough splitter to discourage flow migration but not long enough that the flow and pressure communication between core and bypass is affected which may adversely affect the operating range of the fan  42 . In  FIG.  3   , the inlet guide vane  15  also employs an inlet guide vane splitter  24 . Unlike the blade splitter  26  and splitter  25 , the inlet guide vane splitter  24  positioned upstream from the fan blade  42  at the bottom of the inlet guide vane  15  remains axially displaced from the blade splitter  26  to preserve flow communication.  FIG.  4    illustrates an embodiment without an upstream splitter. 
     The leading edge of the blade splitter  26  as shown in  FIG.  3    is located axially just forward of the midchord line  43 , however, it is envisioned that an axial location between ¾ and ½ of the local cord from the trailing edge will obtain the desired balance between stream separation and flow communication.  FIG.  3    illustrates the embodiment in which the leading edge of the blade splitter  26  is located at the ⅔ of the local chord from the fan&#39;s trailing edge. The trailing edge of the blade splitter  26  terminates proximate to the trailing edge of the fan  42  at the interface with the downstream splitter  25 . As noted above the interface may be a seal or discourager  37 . The seal or discourager  37  may be carbon seal, a labyrinth seal, lip seal or another conventional type seal. Favorable characteristics of the seal  37  include minimal interference with the bypass and core flows, minimum friction and minimum manufacturing and assembly cost. Moreover, the seal or discourager  37 , need only restrict flow from the core to the bypass duct, a hundred percent seal is not required. 
       FIG.  5    is an illustration of an additional splitter with multiple fan stages according to embodiments of the disclosed subject matter. The forward fan  42  and rear fan  41  may be nested with a midstream splitter  25   a  between them. In such case, the midstream splitter  25   a  downstream from the inlet guide vane splitter  24  by communication gap  55 , would interface with the blade splitter  26   a  with a seal or discourager  37   a  and terminate prior to the second fan  41  as to preserve a second communication gap  54 , a second blade splitter  26   b , would likewise interface with the second splitter  25   b . An additional guide vane  15   b  may also be between the forward fan  42  and rear fan  41 , intersecting the midstream splitter  25   a . The guide vane  15   b  while shown operating on both the bypass flow  30  and core flow  28 , may also be limited to only one of the flows, likewise the guide vane  15   b  may fixed as shown or adjustable. Thus communication between the streams is maintained while separating the flows the allowing a wide operating range with reduced leakage. 
       FIG.  6    is a detailed illustration of the blade splitter  26  on fan blade  42 . The fan  42  has a leading edge  141 , trailing edge midchord line  43  and a midspan chord  146 . The blade splitter  26  includes a leading edge  126  and trailing edge  127 . The trailing edge  127  interfaces with the leading edge  128  of the downstream splitter  25  via a seal or discourager  37 . The downstream splitter  25  is fixed with respect to the engine casing (not shown). An upstream splitter  24  is axially forward of the fan  42 . As shown in  FIG.  6   , the blade is generally divided radially into thirds, the first third  101  near the root, the middle third  103  and outer third  105 . The blade splitter is preferably located in the middle third  103 . The leading edge  126  of the blade splitter  26  preferably is forward of the midchord  43  and is proximate the midspan cord  146 , the overlap of the blade splitter  26  on the blade being shown as S b  and the length of the midspan chord shown as C local . The ratio of S b /C local  being from ⅞ to ½, preferably from ¾ to ½, and specifically around ⅔rds. 
     The communication gap  55  by which communications between the bypass flow and core flow is maintained is function of the axial distance from the upstream splitter  24  and the leading edge  126  of the blade splitter  26 . The communication gap  55  includes an axial component (A S ) between the trailing edge of the upstream splitter  24  and the leading edge  141  of the fan  42  (A S  is typically minimized, but for the now recognized advantageous communication between flows) and an axial component (A B ) between the leading edge  141  of the fan  42  and the leading edge  126  of the blade splitter  26 . The communication gap (G) equaling A B +A S , (i.e. G is a function of A S  and C local ) where A S  is preferably less than or equal to A B  and non-zero when the overlap is ⅔ or lower. The communication gap  55  may also be less than or equal to the chord length C local  and preferably less than or equal to the overlap S b . For example, where S B  is ½ C local , the gap G may approach ½ C local  with A S  approaching zero, whereas when S B  is ⅞ C local , the gap may be ½ C local , where A S  is greater than A B . The communication gap ranging between ⅛ C local  and C local , preferably between ⅛ C local  and ½ C local . A balance exists between advantageously increasing S B  to minimize leakage while maintaining an adequate communication gap G as to not detrimentally restrict the operating range. 
     The blade splitter may, advantageously, also minimize vibration and dynamics. Typically, shrouds used for this purpose are at higher spans, but while the disclosed shroud is not primarily a vibration reduction feature, but given its structure it may be beneficial to address these issues as well as the aerodynamic and performance discussed herein. 
     While preferred embodiments of the present invention have been described, it is to be understood that the embodiments described are illustrative only and that the scope of the invention is to be defined solely by the appended claims when accorded a full range of equivalence. Many variations and modifications naturally occurring to those of skill in the art from a perusal hereof.