(12)
`
`United States Patent
`Decker et a].
`
`(10) Patent N0.:
`(45) Date of Patent:
`
`US 7,374,403 B2
`May 20, 2008
`
`US007374403B2
`
`( * ) Notice:
`
`(54) LOW SOLIDITY TURBOFAN
`(75) Inventors: g)I§1IE[.Jl;;edP]t)ecl§.r,hLibergy Towréxslhip,
`;
`e er 1c 0 as zucs,
`es
`Chester’ OH (Us); William Joseph
`Solomon Cincinnati OH
`Virginia Louise Wilson, Walton, KY
`(US)
`_
`_
`(73) Ass1gnee: General Electric Company,
`Schenectady’ NY (Us)
`Subject to any disclaimer, the term of this
`patent is extended or adjusted under 35
`U.S.C. 154(b) by 524 days.
`(21) Appl- No: 11/100,752
`(22) Filed:
`Apr. 7, 2005
`(65)
`Prior Publication Data
`US 2006/0228206 A1
`Oct. 12, 2006
`(51) Int. Cl.
`(2006.01)
`F 01D 25/24
`(52) US. Cl. ........................... .. 416/223 R; 416/223 A;
`416/228; 416/238; 416/242; 416/DIG. 5
`(58) Field of Classi?cation Search .................. .. 415/1,
`415/173.1, 192, 220, 222; 416/223 R, 223 A,
`416/228’ 238, 242, 243, DIG 5
`See application ?le for complete search history.
`_
`References Clted
`US. PATENT DOCUMENTS
`4,358,246 A 11/1982 Hanson et a1.
`4,971,520 A 11/1990 Van Houten
`5,167,489 A 12/1992 Wadia et a1.
`5,169,288 A 12/1992 Gliebe et a1.
`5,273,400 A 12/1993 Amr
`5,478,199 A 12/1995 Gliebe
`5,584,660 A 12/1996 Carter et a1.
`5,642,985 A
`7/1997 Spear et a1.
`5,735,673 A *
`4/1998 Matheny et a1. ...... .. 416/223 A
`5,769,607 A
`6/1998 Neely et a1.
`
`(56)
`
`5,810,555 A
`5,906,179 A
`6,048,174 A
`6,059,532 A
`6,071,077 A
`6,315,521 B1
`6,328,533 B1
`6,338,609 B1
`6,368,061 B1
`6,386,830 B1
`
`9/1998 Savage et a1.
`5/1999 Capdevila
`40000 Samit et a1‘
`5/2000 Chen et a1.
`6/2000 Rowlands
`11/2001 Hunt
`12/2001 Decker et a1.
`1/2002 Decker et a1.
`4/2002 capdevila
`5/2002 Slipper et a1.
`
`(Continued)
`OTHER PUBLICATIONS
`Cumpsty, “Compressor Aerodynamics,” 1989, pp: Cover, copyr.,
`variables, 214, and 215.
`
`(Continued)
`Primary Examinerilgor Kershteyn
`(74)Allorney, Agent, OrFirmiWilliam S.Andes; Francis L.
`Conte
`(57)
`
`ABSTRACT
`
`A turbofan includes a row of fan blades extending from a
`Supporting disk inside an annular easing Each blade
`includes an airfoil having opposite pressure and suction
`sides extending radially in span between a root and tip and
`axially in chord between leading and trailing edges. Adja
`cent airfoils de?ne corresponding ?ow passages therebe
`tween for pressuriZing air. Each airfoil includes stagger
`increasing between the root and tip, and the ?ow passage has
`a mouth between the airfoil leading edge and the suction side
`of an adjacent airfoil and converges to a throat aft from the
`mouth. The row includes no more than twenty fan blades
`having low tip solidity for increasing the width of the
`passage throat.
`
`28 Claims, 6 Drawing Sheets
`
`GE-1007.001
`
`

`
`US 7,374,403 B2
`Page 2
`
`US. PATENT DOCUMENTS
`6,471,474 B1* 10/2002 Mielke et al. ......... .. 415/199.4
`6,524,070 B1* 2/2003
`416/193 A
`RE38,040 E
`3/2003
`6,561,760 B2
`5/2003
`6,561,761 B1
`5/2003 Decker et al.
`6,562,227 B2
`5/2003 Lamphere et al.
`6,991,428 B2* 1/2006 Crane .......................... .. 416/2
`OTHER PUBLICATIONS
`Kandebo, “Geared-Turbofan Engine Design Targets Cost, Com
`pleXity,” Av. Week & Space Tech., vol. 148, No. 8, Feb. 1998, 2
`pages.
`www.rolls-royce.com, “Trent 1000,” copyright 2004, 3 pages.
`
`A. Wadia et al, “Forward Swept Rotor Studies in Multistage Fans
`with Inlet Distortion,” ASME Turbo EXpo 2002, Amsterdam, The
`Netherlands, Jun. 2002, pp. 1-11.
`Pratt & Whitney Canada, “PW 500,” www.pwc.ca, copyright 2000
`2003, 2 pages.
`Aerospace Engineering Online, “Pratt & Whitney’s NeXt Leap in
`Engine Technologies,” www.sae.org, downloaded Feb. 18, 2005, 3
`pages.
`Aerospace Engineering Online, “Pratt & Whitney Gearing up the
`PW 800,” www.sae.org, Aug. 2001, 4 pages.
`Kandebo, “Military Technologies Finging Homes in Commercial
`Engines,” www.AviationNow.com, Jun. 1999, 6 pages.
`* cited by examiner
`
`GE-1007.002
`
`

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`U.S. Patent
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`May 20, 2008
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`Sheet 1 0f 6
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`US 7,374,403 B2
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`GE-1007.003
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`

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`U.S. Patent
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`May 20, 2008
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`Sheet 2 0f 6
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`US 7,374,403 B2
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`GE-1007.004
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`

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`GE-1007.005
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`

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`U.S. Patent
`
`May 20, 2008
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`Sheet 4 0f 6
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`US 7,374,403 B2
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`LOADING>O.29
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`FIG. 4
`
`GE-1007.006
`
`

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`U.S. Patent
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`May 20, 2008
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`Sheet 5 0f 6
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`US 7,374,403 B2
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`GE-1007.007
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`U.S. Patent
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`May 20, 2008
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`Sheet 6 0f 6
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`US 7,374,403 B2
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`Lwm
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`w?
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`GE-1007.008
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`

`
`1
`LOW SOLIDITY TURBOFAN
`
`US 7,374,403 B2
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`BACKGROUND OF THE INVENTION
`
`The present invention relates generally to gas turbine
`engines, and, more speci?cally, to turbofan aircraft engines.
`In a turbofan engine air is pressurized in a compressor and
`mixed With fuel in a combustor for generating hot combus
`tion gases. A high pressure turbine (HPT) extracts energy
`from the combustion gases to poWer the compressor. A loW
`pressure turbine (LPT) extracts additional energy from the
`combustion gases to poWer the fan disposed upstream from
`the compressor.
`The primary design objective of aircraft turbofan engines
`is to maximize ef?ciency thereof for propelling an aircraft in
`?ight, and correspondingly reduce fuel consumption.
`Accordingly, the various cold and hot section rotor and
`stator components Which de?ne the internal ?oW passages
`for the pressurized air and combustion gases, and Which
`extract energy from those gases, are speci?cally designed for
`maximizing the ef?ciency thereof While correspondingly
`obtaining a long useful life.
`The turbofan itself includes a roW of large fan rotor blades
`extending radially outWardly from the perimeter of a sup
`porting rotor disk. The fan is poWered by the LPT for
`pressurizing the incident air for producing a majority of
`propulsion thrust discharged from the fan outlet. Some of the
`fan air is channeled into the compressor Wherein it is
`pressurized and mixed With fuel for generating the hot
`combustion gases from Which energy is extracted in the
`various turbine stages, and then discharged through a sepa
`rate core engine outlet.
`Turbofan engines are continually being developed and
`improved for maximizing their thrust capability With the
`greatest aerodynamic ef?ciency possible. Since the fan pro
`duces a substantial amount of thrust during operation, noise
`is also generated therefrom and should be reduced as much
`as possible consistent With the various competing design
`objectives.
`For example, fan blades are typically designed for maxi
`mizing the aerodynamic loading thereof to correspondingly
`maximize the amount of propulsion thrust generated during
`operation. HoWever, fan loading is limited by stall, ?utter, or
`other instability parameters of the air being pressurized.
`Accordingly, modern turbofan engines are designed With
`a suitable value of stability and stall margin over their
`operating cycle from takeolf to cruise to landing of the
`aircraft to ensure acceptable operation and performance of
`the engine Without overloading the capability of the turbo
`fan.
`Furthermore, modern turbofan engines have relatively
`large diameter turbofans Which rotate at suf?cient rotary
`velocity to create supersonic velocity of the blade tips
`relative to the incident air stream. The blade tips are there
`fore subject to the generation of shock Waves as the air is
`channeled and pressurized in the corresponding ?oW pas
`sages de?ned betWeen adjacent fan blades.
`Accordingly, each fan blade is speci?cally tailored and
`designed from its radially inner platform to its radially outer
`tip and along its circumferentially opposite pressure and
`suction sides Which extend in chord axially betWeen the
`opposite leading and trailing edges thereof. The pressure
`side of one airfoil de?nes With the suction side of an adjacent
`airfoil the corresponding ?oW passage from root to tip of the
`blades through Which the air is channeled during operation.
`
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`Each airfoil is typically tWisted With a corresponding
`angle of stagger from root to tip, With airfoil tips being
`aligned obliquely betWeen the axial and circumferential
`directions of the fan.
`During operation, the incoming ambient air ?oWs at
`different relative velocities through the inter-blade ?oW
`passages from root to tip of the blades including subsonic
`air?oW at the blade roots and radially outWardly thereof up
`to the supersonic velocity of the air at the blade tips in
`various portions of the operating range.
`Fan stall margin is a fundamental design requirement for
`the turbofan and is affected by the aerodynamic fan loading,
`the fan solidity, and the fan blade aspect ratio. These are
`conventional parameters, With the fan loading being the rise
`in speci?c enthalpy across the fan blades divided by the
`square of the tip speed.
`Blade solidity is the ratio of the blade chord, represented
`by its length, over the blade pitch, Which is the circumfer
`ential spacing of the blades at a given radius or diameter
`from the axial centerline axis. In other Words, blade pitch is
`the circumferential length at a given diameter divided by the
`number of blades in the full fan blade roW. And, the fan blade
`aspect ratio is the radial height or span of the airfoil portion
`of the blade divided by its maximum chord.
`Conventional experience or teachings in the art indicate
`that When inlet Mach numbers are su?iciently high that
`passage shock can separate the suction surface boundary
`layer of the air in the inter-blade ?oW passages, good
`ef?ciency requires that the solidity should be high to alloW
`the How to reattach.
`In one exemplary or reference turbofan found in public
`use and on sale for more than a year in the USA, a large
`diameter turbofan having tWenty-tWo fan blades in the full
`roW has a relatively high solidity at the blade tips of about
`1.29. These fan blades are used in a high bypass ratio
`turbofan engine With a bypass ratio over 7, With the corre
`sponding pressure ratio over the fan blades being relatively
`high in value and greater than about 1.5. The large fan
`diameter e?fects supersonic velocity of the blade tips during
`operation Which correspondingly generates normal shock
`Waves at the airfoil tips during operation which affect
`performance.
`Conventional design practice for turbofan ef?ciency and
`adequate fan stall margin typically require the relatively
`high tip solidity Which is generally equal to the fan tip
`relative Mach number at the design point, such as cruise
`operation. In other Words, the tip Mach number is suitably
`greater than one (1.0) for supersonic How, and the fan tip
`solidity is correspondingly greater than one and generally
`equal to the tip relative Mach number for good designs.
`The design considerations disclosed above are merely
`some of the many competing design parameters in designing
`a modern turbofan primarily for good aerodynamic perfor
`mance and ef?ciency, as Well as for good mechanical
`strength for ensuring a long useful life thereof. Each fan
`blade tWists from root to tip, and the opposite pressure and
`suction sides thereof also vary in con?guration to speci?
`cally tailor the How passages from root to tip for maximizing
`fan ef?ciency With suitable stall margin and mechanical
`strength.
`The resulting turbofan design is a highly complex design
`With three dimensional variation of the pressure and suction
`sides of the individual airfoils across their axial chord and
`over their radial span. And, the individual fan blades coop
`erate With each other in the full roW of blades to de?ne the
`inter-blade ?oW passages and to effect the resulting aerody
`namic performance and stall margin of the entire fan.
`
`GE-1007.009
`
`

`
`3
`Accordingly, it is desired to further improve the e?iciency
`of the modern turbofan While maintaining adequate stability
`and stall margin notwithstanding the various competing
`design objectives addressed in part above.
`BRIEF DESCRIPTION OF THE INVENTION
`
`Aturbofan includes a roW of fan blades extending from a
`supporting disk inside an annular casing. Each blade
`includes an airfoil having opposite pressure and suction
`sides extending radially in span betWeen a root and tip and
`axially in chord betWeen leading and trailing edges. Adja
`cent airfoils de?ne corresponding ?oW passages therebe
`tWeen for pressurizing air. Each airfoil includes stagger
`increasing betWeen the root and tip, and the ?oW passage has
`a mouth betWeen the airfoil leading edge and the suction side
`of an adjacent airfoil and converges to a throat aft from the
`mouth. The roW includes no more than tWenty fan blades
`having loW tip solidity for increasing the Width of the
`passage throat.
`BRIEF DESCRIPTION OF THE DRAWINGS
`The invention, in accordance With preferred and exem
`plary embodiments, together With further objects and advan
`tages thereof, is more particularly described in the folloWing
`detailed description taken in conjunction With the accom
`panying draWings in Which:
`FIG. 1 is a partly schematic isometric vieW of a turbofan
`in an aircraft engine for poWering an aircraft in ?ight.
`FIG. 2 is an axial sectional vieW through the turbofan
`portion of the engine illustrated in FIG. 1 and taken along
`line 2-2.
`FIG. 3 is a forWard-facing-aft elevational vieW of the
`turbofan illustrated in FIG. 1 and taken along line 3-3.
`FIG. 4 is a top planiform vieW of tWo adjacent fan blades
`illustrated in FIG. 3 and taken generally along line 4-4 in
`conjunction With a corresponding ?owchart.
`FIG. 5 is a forWard-facing-aft elevational vieW, like FIG.
`3, of a turbofan in accordance With another embodiment.
`FIG. 6 is a top planiform vieW of tWo adjacent fan blades
`in the turbofan illustrated in FIG. 5 and taken along line 6-6.
`DETAILED DESCRIPTION OF THE
`INVENTION
`
`Illustrated in FIG. 1 is a gas turbine engine 10 con?gured
`for poWering an aircraft 12 in ?ight, and suitably mounted
`therein. The engine is axisymmetrical about a longitudinal or
`axial centerline axis and includes a fan or turbofan 14
`suitably mounted coaxially inside a surrounding annular fan
`casing 16.
`During operation, ambient air 18 enters the inlet end of
`the fan 14 and is pressurized thereby for producing propul
`sion thrust for propelling the aircraft in ?ight. A portion of
`the fan air is suitably channeled in turn through a loW
`pressure or booster compressor 20 and a high pressure
`compressor 22 Which further pressurize the air in turn.
`The pressurized air is mixed With fuel in an annular
`combustor 24 for generating hot combustion gases 26 Which
`are discharged in the doWnstream direction. A high pressure
`turbine (HPT) 28 ?rst receives the hot gases from the
`combustor for extracting energy therefrom, and is folloWed
`in turn by a loW pressure turbine (LPT) 30 Which extracts
`additional energy from the combustion gases discharged
`from the HPT. The HPT is joined by one shaft or rotor to the
`high pressure compressor 22, and the LPT is joined by
`
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`4
`another shaft or rotor to both the booster compressor 20 and
`the fan 14 for poWering thereof during operation.
`The exemplary turbofan engine 10 illustrated in FIG. 1
`may have any conventional con?guration and operation for
`poWering an aircraft in ?ight from takeolf to cruise to
`landing, but is modi?ed as further described hereinbeloW for
`increasing the aerodynamic e?iciency of the fan 14 While
`maintaining suitable stability and stall margin thereof during
`the operating cycle.
`More speci?cally, FIGS. 1 and 2 illustrate an exemplary
`embodiment of the turbofan 14 Which includes a roW of fan
`rotor blades 32 extending radially outWardly in span from
`the perimeter rim of a supporting rotor disk 34. As shoWn in
`FIG. 2, each blade includes an airfoil 36 extending out
`Wardly from a platform 38 de?ning the radially inner
`boundary of the fan air ?oWpath, Which platform may be
`integrally formed With the airfoil or a separate component.
`Each blade also includes an integral dovetail 40 extending
`radially inWardly from the airfoil beloW the platform for
`mounting each blade in a corresponding dovetail slot in the
`rim of the rotor disk.
`The fan blades may be made from suitable high strength
`materials like titanium or carbon ?ber composites. For
`example, the majority of the blade may be formed of carbon
`?ber composite reinforced With titanium shields along the
`leading and trailing edges, and along the tip.
`As illustrated in FIGS. 1 and 2, each airfoil 36 has a
`suitable aerodynamic con?guration including a generally
`concave pressure side 42 and a circumferentially opposite,
`generally convex suction side 44. The opposite sides of each
`airfoil extend radially in span from the inner root end thereof
`at the platform 38 to the radially outer distal tip 46 disposed
`closely adjacent to the fan stator casing 16 for providing a
`relatively small tip clearance or gap therebetWeen.
`As shoWn in FIGS. 2 and 3, each airfoil extends axially in
`chord C betWeen opposite leading and trailing edges 48,50,
`With the chord varying in length over the span of the airfoil.
`As shoWn in FIG. 4, adjacent airfoils 36 de?ne circum
`ferentially therebetWeen corresponding ?oW passages 52 for
`pressurizing the air 18 during operation. Each of the airfoils
`36 includes stagger or tWist represented by the stagger angle
`A from the axial or longitudinal axis, Which stagger
`increases betWeen the root and tip of the airfoil.
`For example, the stagger angle A at the blade tip may be
`substantial, and about 60 degrees, to position the leading
`edge 48 of one airfoil circumferentially adjacent but axially
`spaced from the suction side 44 of the next adjacent airfoil
`aft from the leading edge thereof to de?ne a corresponding
`mouth 54 for the ?oW passage betWeen the opposing pres
`sure and suction sides of the adjacent airfoils. The contours
`and stagger of the adjacent airfoils over the radial span of the
`blades cause each ?oW passage to converge or decrease in
`?oW area to a throat 56 of minimum ?oW area spaced aft
`from the mouth along most, if not all, of the radial span.
`As further illustrated in FIG. 4, the relatively high airfoil
`staggerA also positions the trailing edge 50 of one airfoil 36
`circumferentially adjacent to the pressure side 42 of the next
`adjacent airfoil While also being spaced axially therefrom in
`the tip region to de?ne a corresponding discharge or outlet
`58 for the corresponding ?oW passage betWeen adjacent
`airfoils. In this Way, the incoming air 18 is channeled in the
`corresponding ?oW passages 52 betWeen adjacent airfoils as
`they rotate clockWise in FIGS. 1,3, and 4 for pressurizing the
`air to produce the propulsion thrust during operation.
`FIGS. 1-4 illustrate in general the typical con?guration of
`a modern turbofan aircraft engine having a roW of fan blades
`With corresponding stagger or tWist from root to tip. As
`
`GE-1007.010
`
`

`
`5
`indicated in the Background section, there are many com
`peting design parameters for the turbofan for balancing fan
`ef?ciency With stability and stall margin and With aero
`mechanical parameters affecting ?utter and noise and With
`mechanical strength of the fan blade subject both to cen
`trifugal force during operation and aerodynamic loading.
`FIG. 4 illustrates schematically a method of improving
`aerodynamic ef?ciency of the turbofan engine 10 illustrated
`in FIG. 1 by derivation for example. Modern turbofan
`engines are typically derived from existing engines having
`proven experience in commercial service. Corresponding
`changes or modi?cations thereof may then be effected in
`accordance With conventional design practices, Which, hoW
`ever, must be balanced in vieW of the various competing
`parameters such as ef?ciency and stall margin, for example.
`Further increasing e?iciency and aerodynamic loading typi
`cally requires reduction in stall margin, and must therefore
`be balanced for overall performance.
`FIG. 4 illustrates schematically a pre-existing or conven
`tional design of a fan 60 for use in the type of turbofan
`engine illustrated in FIG. 1. This pre-existing fan has a full
`complement of only tWenty-tWo fan blades of suitably large
`outer diameter D for effecting supersonic air?oW at the tips
`during operation.
`The pre-existing fan 60 also has a corresponding solidity
`Which is a conventional parameter equal to the ratio of the
`airfoil chord C, as represented by its length, divided by the
`circumferential pitch P or spacing from blade to blade at the
`corresponding span position or radius.
`The circumferential pitch is equal to the circumferential
`length at the speci?c radial span divided by the total number
`of fan blades in the blade roW. Accordingly, the solidity is
`directly proportional to the number of blades and chord
`length and inversely proportional to the diameter as shoWn
`schematically in FIG. 4.
`As indicated above, modern design practice requires the
`solidity of the blades at the airfoil tips to be generally similar
`in magnitude to the relative Mach number of the How stream
`at the airfoil tips. In this exemplary embodiment, the tip
`solidity of the pre-existing fan 60 is relatively high at about
`1.29 and corresponds Well With a similar tip relative Mach
`number of also about 1.29.
`Conventional practice as indicated above requires rela
`tively high tip solidity for maintaining good ef?ciency in a
`supersonic blade tip design subject to shock in the How
`passages betWeen the adjacent airfoils, and therefore
`increasing solidity is one option, of the various design
`parameters for a modern turbofan, in producing a derivative
`fan. Or, tip solidity may remain the same, or equal, in the
`derivative fan.
`HoWever, it has been discovered that notWithstanding this
`conventional practice for relatively high solidity in modern
`turbofans, a substantial improvement in ef?ciency While
`maintaining adequate stability and stall margin may be
`obtained by decreasing tip solidity, and not increasing tip
`solidity. As indicated above, solidity is proportional to the
`number of fan blades and the ratio of the airfoil chord
`divided by the diameter of the fan.
`Accordingly, solidity may be decreased by decreasing the
`number of fan blades, decreasing the airfoil chord, or
`increasing the outer diameter of the fan. HoWever, the fan
`outer diameter is typically a given parameter for a speci?
`cally siZed turbofan engine. And, it has been further discov
`ered that reducing solidity by reducing the length of the
`chord is detrimental to turbofan e?iciency, Whereas reducing
`the blade count to reduce solidity can improve turbofan
`ef?ciency.
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`6
`FIG. 4 illustrates schematically these various options in
`turbofan design based on the blade solidity. Decreasing
`solidity by reducing the chord to diameter C/D ratio main
`tains constant the number of fan blades, for example tWenty
`tWo, yet analysis indicates a reduction in ef?ciency.
`Correspondingly, the chord to diameter C/D ratio may
`remain constant or equal betWeen the turbofan designs, With
`instead the number of fan blades being reduced to tWenty or
`eighteen in the preferred embodiments.
`Accordingly, aerodynamic ef?ciency may be improved in
`the turbofan engine 10 illustrated in FIG. 1 by deriving the
`fan 14 from the pre-existing fan 60 and reducing the solidity
`at the airfoil tips by reducing the number of blades from
`tWenty-tWo to either tWenty or eighteen, for example, While
`maintaining substantially equal or constant the same ratio of
`the tip chord over the tip diameter C/ D in the derived fan 14
`as originally found in the preexisting fan.
`Furthermore, the reduction in number of fan blades
`increases the circumferential pitch P betWeen the airfoils and
`increases the How area of the How passages 52, in particular
`at the throats 56 thereof, for reducing ?oW blockage during
`operation, and speci?cally at the airfoil tips subject to
`supersonic operation.
`Accordingly, the derived turbofan 14 illustrated in FIGS.
`1-4 includes no more than tWenty of the fan blades 32
`effected by reducing the tip solidity Which has a relatively
`loW magnitude at the tips 46 to position the leading edge 48
`of each tip 46 circumferentially near the trailing edge 50 of
`the next adjacent tip 46, and correspondingly increase the
`Width of the throat 54 normal or perpendicular betWeen the
`opposing pressure and suction sides of adjacent airfoils.
`The reduction in fan blade number While maintaining
`substantially constant the chord to diameter C/D ratio at the
`airfoil tips has signi?cant advantages in the neW turbofan
`including an increase in ef?ciency While maintaining
`adequate stability and stall margin, as Well as reducing noise,
`as Well as reducing Weight and cost due to the feWer fan
`blades.
`Quite signi?cant in the loW solidity turbofan design is the
`substantial reduction in How blockage at the passage throats
`Which more than o?fsets the decreased solidity effect on
`aerodynamic performance. Modern computational ?oW
`dynamics analysis predicts that loWer solidity through
`reduced blade number is bene?cial to cruise ef?ciency
`Whereas loWer solidity through reduction of the chord to
`diameter C/ D ratio Would be detrimental to cruise ef?ciency,
`Which has been con?rmed by testing.
`FIG. 3 is a front vieW of the turbofan 14, With FIG. 4
`being a top planiform vieW Which illustrate the substantial
`change in appearance of the turbofan as opposed to typical
`high solidity turbofans in Which the adjacent fan blades
`substantially overlap each other circumferentially due to the
`high solidity and high stagger of the airfoils.
`In contrast, the tip solidity of the turbofan illustrated in
`FIGS. 3 and 4 is relatively loW in magnitude, While still
`being greater than about 1.0 to provide a circumferential gap
`G betWeen the leading and trailing edges 48,50 of adjacent
`tips 46.
`In particular, since the loW solidity is effected by reducing
`the blade count instead of reducing the chord to diameter
`C/D ratio, this ratio, and chord, remain relatively large in
`value, Which along With the increased circumferential pitch
`P and large staggerA of the airfoils is effective to provide the
`circumferential gap G locally betWeen the leading and
`trailing edges of the adjacent tips.
`The con?guration of the How passage 52 illustrated in
`FIG. 4 is particularly important to e?icient operation of the
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`fan, and in particular at the airfoil tips subject to supersonic
`How. The speci?c pro?les of the pressure and suction sides
`of the individual airfoils, the lateral thickness of the airfoil,
`the root to tip stagger A of the airfoils and, of course, the
`reduced solidity due to the reduction in blade count While
`maintaining equal the chord to diameter C/D ratio are all
`used to de?ne each ?oW passage 52.
`In particular, the airfoil tips 36 are locally angled and vary
`in Width betWeen the leading and trailing edges 48,50 to
`typically converge the How passage 52 at the airfoil tips
`from the mouth 54 to the throat 56 and then diverge the How
`passage also at the tip from the throat 56 to the outlet 58.
`Alternatively, the mouth and throat of the How passages at
`the airfoil tips may be coincident in one plane at the leading
`edges, With the How passages still diverging aft from the
`throats at the leading edges to the passage outlets at the
`trailing edges.
`The convergence angle or slope betWeen the mouth and
`the throat, and the divergence angle or slope betWeen the
`throat and the outlet may be speci?cally designed for
`maximizing ef?ciency during supersonic operation of the
`blade tips in Which aerodynamic shock is generated as the
`air?oW is reduced in speed in the converging portion to
`choked How of Mach 1 at the throat 56 folloWed in turn by
`subsonic diffusion in the diverging portion of the How
`passage from or aft of the throat to the outlet.
`The ratio of the How area at the passage outlet 58 over the
`How area at the throat 56 is a conventional measure of
`effective camber of the airfoils. The actual amount of airfoil
`camber at the tips thereof may be increased slightly over a
`conventional turbofan design to alloW the turbofan to toler
`ate the loWer tip solidity during part-speed operation.
`As indicated above, a modern turbofan is designed for an
`operating range from takeoff to cruise to landing, With cruise
`operation being predominant and for Which maximum effi
`ciency and operability are desired. HoWever, part-speed
`performance must also be considered in good turbofan
`design and accommodated by the higher camber introduced
`at the blade tips for the loW solidity turbofan design.
`Accordingly, part-speed operability may be improved by
`increasing the camber of the airfoils 36 at the tips 46 thereof
`in conjunction With the reduction in solidity by reduction in
`blade count.
`Since improved ef?ciency of the fan may be obtained
`through lowering solidity, the turbofan design may itself be
`otherWise conventional except as modi?ed in accordance
`With the present disclosure. For example, the airfoils 36
`illustrated in FIGS. 1-4 are relatively large in diameter for
`supersonic tip operation in a modern turbofan engine With a
`substantial pressure ratio of about 1.5. The corresponding
`bypass ratio of the fan air Which bypasses the core engine is
`about 7.5 or greater. And, the airfoils may be provided With
`suitable aerodynamic sWeep Which is preferably forWard or
`negative (S—) at the tips 46 of the airfoils, and preferably
`negative along both the leading and trailing edges 48,50
`thereof.
`The individual airfoils may have a large chord barreling
`near their midspan as illustrated in FIG. 2 With aft or positive
`aerodynamic sWeep (S+) along a portion of the leading edge
`above the midspan if desired. This form of modem turbofan
`blade is disclosed in substantial detail in Us. Pat. No.
`6,328,533, and is incorporated herein by reference.
`Aerodynamic sWeep is also a conventional term of art and
`is disclosed in detail in Us. Pat. No. 5,167,489, also
`incorporated herein by reference. The forWard tip sWeep in
`the fan blades improves ef?ciency during supersonic opera
`tion of the blade tips.
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`FIGS. 1 and 2 also illustrate that the turbofan includes an
`annular tip shroud 62 suitably mounted ?ush inside the fan
`stator casing 16 and directly surrounding the airfoil tips 46
`Which are positioned closely adjacent thereto to de?ne a
`correspondingly small tip clearance thereWith. The tip
`shroud 62 may be conventional in con?guration, such as a
`lightWeight honeycomb structure, With a substantially
`smooth inner surface facing the blade tips. The loW solidity
`turbofan enjoys improved ef?ciency While maintaining
`adequate stability and stall margin Without the need for
`stability enhancing features such as annular grooves Which
`could otherWise be formed in the tip shroud.
`As shoWn in FIG. 2, the fan casing 16 is spaced radially
`outWardly from an inner casing 64 Which surrounds the core
`engine to de?ne an annular bypass duct 66 radially therebe
`tWeen. The aft end of the bypass duct 66 de?nes the outlet
`for a majority of the fan air used in producing propulsion
`thrust for the engine.
`Spaced doWnstream or aft from the roW of fan blades 32
`is a roW of outlet guide vanes 68 extending radially inWardly
`from the fan casing 16 to join the inner casing 64. The
`number of vanes 68 is preferably more than tWice the
`number of the fan blades 32 for reducing noise from the fan
`during operation.
`Noise reduction, and in particular spinning mode noise, is
`disclosed in Us. Pat. No. 5,169,288, incorporated herein by
`reference, Which patent may be used for determining the
`speci?c number of vanes 68 relative to the speci?c number
`of fan blades, and for example may number 48 vanes for
`both the tWenty and eighteen fan blade species.
`FIG. 2 illustrates another feature Which may be intro
`duced into the turbofan. In particular, the airfoil tips 46 may
`have an axially arcuate contour radially outWardly betWeen
`the leading and trailing edges, and the adjacent tip shroud 62
`may have a complementary axially arcuate contour radially
`inWardly for maintaining a substantially uniform tip clear
`ance radially therebetWeen, and axially betWeen the leading
`and trailing edges 48,50 of the airfoils. In one embodiment,
`the forWard portion of the airfoil tip 46 is convex folloWed
`in turn by a concave aft portion. Correspondingly, the tip
`shroud 46 has a forWard concave portion folloWed by a
`convex aft portion for reducing tip losses and How blockage
`during supersonic operation