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`US009121412B2
`
`c12) United States Patent
`Gallagher et al.
`
`(10) Patent No.:
`(45) Date of Patent:
`
`US 9,121,412 B2
`*Sep.1,2015
`
`(54) EFFICIENT, LOW PRESSURE RATIO
`PROPULSOR FOR GAS TURBINE ENGINES
`
`(75)
`
`Inventors: Edward J. Gallagher, West Hartford,
`CT (US); Jun Jiang, Glastonbury, CT
`(US); Becky E. Rose, Colchester, CT
`(US); Jason Elliott, Huntington, IN
`(US); Anthony R. Bifulco, Ellington,
`CT (US)
`
`(73) Assignee: United Technologies Corporation,
`Hartford, CT (US)
`
`( *) Notice:
`
`Subject to any disclaimer, the term ofthis
`patent is extended or adjusted under 35
`U.S.C. 154(b) by 997 days.
`
`This patent is subject to a terminal dis(cid:173)
`claimer.
`
`(21) Appl. No.: 13/176,255
`
`(22) Filed:
`
`Jul. 5, 2011
`
`(65)
`
`(51)
`
`(52)
`
`(58)
`
`Prior Publication Data
`
`US 2013/0008144 Al
`
`Jan. 10,2013
`
`(2006.01)
`(2006.01)
`(2006.01)
`(2006.01)
`(2006.01)
`(2006.01)
`
`Int. Cl.
`F04D29/32
`F02K 3106
`F02K 3102
`F04D25/02
`F02C 7136
`F02K 31075
`U.S. Cl.
`CPC ................ F04D 291321 (2013.01); F02C 7136
`(2013.01); F02K 3106 (2013.01); F04D 251022
`(2013.01); F04D 251024 (2013.01); F02K
`31075 (2013.01); F05B 22201306 (2013.01);
`F05D 226014031 (2013.01)
`Field of Classification Search
`CPC combination set(s) only.
`See application file for complete search history.
`
`(56)
`
`References Cited
`
`U.S. PATENT DOCUMENTS
`
`3,468,473 A * 9/1969 Chilman et al.
`.............. 415/129
`3,747,343 A * 7/1973 Rosen .......................... 60/226.l
`4,569,199 A * 2/1986 Klees et al. .................. 60/226.l
`4,732,532 A * 3/1988 Schwaller et al. ............ 415/119
`5,769,607 A * 6/1998 Neely et al. ................... 416/189
`6,004,095 A * 12/1999 Waitz et al. ................... 415/119
`6,058,694 A *
`512000 Ackerman et al. ........... 60/39.08
`6,059,532 A *
`512000 Chen et al. ................ 416/223 A
`6,195,983 Bl
`3/2001 Wadiaet al.
`6,709,239 B2
`3/2004 Chandraker
`7,107,756 B2
`912006 Rolt
`7,175,393 B2 *
`212007 Chandraker .............. 416/223 A
`7,374,403 B2 * 5/2008 Decker et al. ............. 416/223 R
`7,758,306 B2
`712010 Burton et al.
`7,770,377 B2
`8/2010 Rolt
`8,667,775 Bl* 3/2014 Kisska et al. ................ 60/226.l
`2008/0095633 Al
`4/2008 Wilson
`2008/0226454 Al* 9/2008 Decker et al. . ................ 416/203
`2010/0089019 Al
`4/2010 Knight et al.
`2010/0162683 Al*
`712010 Grabowski et al. .......... 601226.3
`2010/0260609 Al
`10/2010 Wood et al.
`2013/0008146 Al*
`1/2013 Gallagher et al.
`........... 601226.3
`2013/0014488 Al*
`1/2013 Gallagher et al. ........... 60/226.l
`* cited by examiner
`
`Primary Examiner - Phutthiwat Wongwian
`Assistant Examiner - Rene Ford
`(74) Attorney, Agent, or Firm - Carlson, Gaskey & Olds,
`P.C.
`
`(57)
`
`ABSTRACT
`
`A gas turbine engine includes a spool, a turbine coupled to
`drive the spool and a propulsor that is coupled to be driven by
`the turbine through the spool. A gear assembly is coupled
`between the propulsor and the spool such that rotation of the
`spool drives the propulsor at a different speed than the spool.
`The propulsor includes a hub and a row of propulsor blades
`that extends from the hub. The row includes no more than 16
`of the propulsor blades.
`
`11 Claims, 2 Drawing Sheets
`
`GE-1001.001
`
`

`
`U.S. Patent
`
`Sep.1,2015
`
`Sheet 1of2
`
`US 9,121,412 B2
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`N
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`GE-1001.002
`
`

`
`U.S. Patent
`
`Sep.1,2015
`
`Sheet 2of2
`
`US 9,121,412 B2
`
`42
`
`FIG.2
`
`GE-1001.003
`
`

`
`US 9,121,412 B2
`
`1
`EFFICIENT, LOW PRESSURE RATIO
`PROPULSOR FOR GAS TURBINE ENGINES
`
`STATEMENT REGARDING FEDERALLY
`SPONSORED RESEARCH OR DEVELOPMENT
`
`This invention was made with government support under
`contract number DTFAWA-10-C-00041 awarded by United
`States FederalAviationAdministration ("FAA"). The govern-
`ment has certain rights in the invention.
`
`BACKGROUND
`
`2
`the following detailed description. The drawings that accom(cid:173)
`pany the detailed description can be briefly described as fol(cid:173)
`lows.
`FIG. 1 is a schematic cross-section of a gas turbine engine.
`FIG. 2 is a perspective view ofa fan section of the engine of
`FIG. 1.
`
`DETAILED DESCRIPTION OF THE PREFERRED
`EMBODIMENT
`
`10
`
`FIG. 1 schematically illustrates a gas turbine engine 20.
`The gas turbine engine 20 may be a two-spool turbofan that
`generally incorporates a fan section 22, a compressor section
`24, a combustor section 26 and a turbine section 28. Altema-
`15 tive engine architectures may include a single-spool design, a
`three-spool design, or an open rotor design rather than the
`ducted design that is shown, among other systems or features.
`The fan section 22 drives air along a bypass flow passage B
`while the compressor section 24 drives air along a core flow
`20 passage C for compression and communication into the com(cid:173)
`bustor section 26. Although depicted as a turbofan gas turbine
`engine, it is to be understood that the concepts described
`herein are not limited to use with turbofans and the teachings
`may be applied to other types of gas turbine engines.
`The engine 20 includes a low speed spool 30 and high
`speed spool 32 mounted for rotation about an engine central
`longitudinal axis A relative to an engine static structure 36 via
`several bearing systems 38. The low speed spool 30 generally
`includes an inner shaft 40 that is coupled with a propulsor 42,
`30 a low pressure compressor 44 and a low pressure turbine 46.
`The low pressure turbine 46 drives the propulsor 42 through
`the inner shaft 40 and a gear assembly 48, which allows the
`low speed spool 30 to drive the propulsor 42 at a different (e.g.
`lower) angular speed.
`An exemplary gas turbine engine includes a turbine 35
`The high speed spool 32 includes an outer shaft 50 that is
`coupled with a high pressure compressor 52 and a high pres-
`coupled through a spool to drive a propulsor. A gear assembly
`sure turbine 54. A combustor 56 is arranged between the high
`is coupled between the propulsor and the spool such that
`pressure compressor 52 and the high pressure turbine 54. The
`rotation of the spool drives the propulsor at a different speed
`than the spool. The propulsor includes a hub and a row of
`inner shaft 40 and the outer shaft 50 are concentric and rotate
`propulsor blades that extend from the hub. The row includes 40 about the engine central longitudinal axis A, which is col-
`no more than 16 propulsor blades.
`linear with their longitudinal axes.
`Core airflow C is compressed by the low pressure compres-
`In another exemplary aspect, a gas turbine engine includes
`a core flow passage and a bypass flow passage defining an
`sor 44 then the high pressure compressor 52, mixed with the
`inlet and an outlet. A propulsor is arranged at the inlet of the
`fuel and burned in the combustor 56, and then expanded over
`bypass flow passage. The propulsor includes a hub and a row 45 the high pressure turbine 54 and low pressure turbine 46. The
`of propulsor blades that extend from the hub. The row
`turbines 54, 46 rotationally drive the respective low speed
`includes no more than 16 of the propulsor blades. The bypass
`spool 30 and high speed spool 32 in response to the expan-
`flow passage has a design pressure ratio of approximately
`s10n.
`1.1-1.35 with regard to inlet pressure and outlet pressure of
`As shown, the propulsor42 is arranged at an inlet 60 of the
`the bypass flow passage.
`50 bypass flow passage B. Air flow through the bypass air pas-
`An exemplary propulsor for use in a gas turbine engine
`sage B exits the engine 20 through an outlet 62 or nozzle. For
`includes a rotor having a row of propulsor blades that extends
`a given design of the propulsor 42, the inlet 60 and the outlet
`radially outwardly from a hub. Each of the propulsor blades
`62, the engine 20 define a design pressure ratio with regard to
`extends radially between a root and a tip and in a chord
`an inlet pressure at the inlet 60 and an outlet pressure at the
`direction between a leading edge and a trailing edge to define 55 outlet 62 of the bypass flow passage B. As an example, the
`a chord dimension. The row of propulsor blades defines a
`design pressure ratio may be determined based upon the
`circumferential pitch with regard to the tips. The row of
`stagnation inlet pressure and the stagnation outlet pressure at
`a design rotational speed of the engine 20. In that regard, the
`propulsor blades has a solidity value defined as the chord
`dimension at the tip divided by the circumferential pitch. The
`engine 20 may optionally include a variable area nozzle 64
`row also includes a number of the propulsor blades that is no 60 within the bypass flow passage B. The variable area nozzle 64
`greater than 16 such that a ratio of the number of propulsor
`is operative to change a cross-sectional area 66 of the outlet 62
`blades to the solidity value is between 8 and 28.
`to thereby control the pressure ratio via changing pressure
`within the bypass flow passage B. The design pressure ratio
`may be defined with the variable area nozzle 64 fully open or
`65 fully closed.
`Referring to FIG. 2, the propulsor 42, which in this
`example is a fan, includes a rotor 70 having a row 72 of
`
`This disclosure relates to gas turbine engines and, more
`particularly, to an engine having a geared turbo fan architec(cid:173)
`ture that is designed to efficiently operate with a high bypass
`ratio and a low pressure ratio.
`The propulsive efficiency of a gas turbine engine depends
`on many different factors, such as the design of the engine and
`the resulting performance debits on the fan that propels the
`engine. As an example, the fan rotates at a high rate of speed
`such that air passes over the blades at transonic or supersonic
`speeds. The fast-moving air creates flow discontinuities or
`shocks that result in irreversible propulsive losses. Addition- 25
`ally, physical interaction between the fan and the air causes
`downstream turbulence and further losses. Although some
`basic principles behind such losses are understood, identify(cid:173)
`ing and changing appropriate design factors to reduce such
`losses for a given engine architecture has proven to be a
`complex and elusive task.
`
`SUMMARY
`
`BRIEF DESCRIPTION OF THE DRAWINGS
`
`The various features and advantages of the disclosed
`examples will become apparent to those skilled in the art from
`
`GE-1001.004
`
`

`
`US 9,121,412 B2
`
`4
`TABLE-continued
`
`Number of Blades. Solidity and Ratio N/R
`
`Number of Blades (N)
`
`Solidity
`
`Ratio N/R
`
`12
`10
`
`0.6
`0.6
`
`20.00
`16.67
`
`The disclosed ratios of N/R enhance the propulsive effi-
`10 ciency of the disclosed engine 20. For instance, the disclosed
`ratios of N/R are designed for the geared turbo fan architec(cid:173)
`ture of the engine 20 that utilizes the gear assembly 48. That
`is, the gear assembly 48 allows the propulsor 42 to rotate at a
`different, lower speed than the low speed spool 30. In com(cid:173)
`bination with the variable area nozzle 64, the propulsor 42 can
`be designed with a large diameter and rotate at a relatively
`slow speed with regard to the low speed spool 30. A relatively
`low speed, relatively large diameter, and the geometry that
`permits the disclosed ratios ofN/R contribute to the reduction
`of performance debits, such as by lowering the speed of the air
`or fluid that passes over the propulsor blades 74.
`Although a combination of features is shown in the illus(cid:173)
`trated examples, not all of them need to be combined to
`realize the benefits of various embodiments of this disclosure.
`In other words, a system designed according to an embodi(cid:173)
`ment of this disclosure will not necessarily include all of the
`features shown in any one of the Figures or all of the portions
`schematically shown in the Figures. Moreover, selected fea(cid:173)
`tures of one example embodiment may be combined with
`selected features of other example embodiments.
`The preceding description is exemplary rather than limit(cid:173)
`ing in nature. Variations and modifications to the disclosed
`examples may become apparent to those skilled in the art that
`do not necessarily depart from the essence of this disclosure.
`The scope oflegal protection given to this disclosure can only
`be determined by studying the following claims.
`
`15
`
`3
`propulsor blades 7 4 that extend a circumferentially around a
`hub 76. Each of the propulsor blades 74 extends radially
`outwardly from the hub 76 between a root 78 and a tip 80 and
`in a chord direction (axially and circumferentially) between a
`leading edge 82 and a trailing edge 84. A chord dimension 5
`(CD) is a length that extends between the leading edge 82 and
`the trailing edge 84 at the tip of each propulsor blade 7 4. The
`row 72 of propulsor blades 7 4 also defines a circumferential
`pitch (CP) that is equivalent to the arc distance between the
`tips 80 of neighboring propulsor blades 74.
`As will be described, the example propulsor 42 includes a
`number (N) of the propulsor blades 74 and a geometry that, in
`combination with the architecture of the engine 20, provides
`enhanced propulsive efficiency by reducing performance
`debits of the propulsor 42.
`In the illustrated example, the number N of propulsor
`blades in the row 72 is no more than 16. In one example, the
`propulsor 42 includes 16 of the propulsor blades 7 4 uniformly
`circumferentially arranged about the hub 76. In other embodi(cid:173)
`ments, the number N is from 10 to 16 and may be any of 11, 20
`12, 13, 14 or 15 blades.
`The propulsor blades 7 4 define a solidity value with regard
`to the chord dimension CD and the circumferential pitch CP.
`The solidity value is defined as a ratio (R) ofCD/CP (i.e., CD
`divided by CP). In embodiments, the solidity value of the 25
`propulsor 42 is between 0.6 and 1.1.
`Additionally, in combination with the given example solid-
`ity values, the engine 20 may be designed with a particular
`design pressure ratio. In embodiments, the design pressure
`ratio may be between 1.1 and 1.35. In a further embodiment 30
`the design pressure ratio may be between 1.2 and 1.3.
`'
`The engine 20 may also be designed with a particular
`bypass ratio with regard to the amount of air that passes
`through the bypass flow passage B and the amount of air that
`passes through the core flow passage C. As an example, the 35
`design bypass ratio of the engine 20 may nominally be 18.
`The propulsor 42 also defines a ratio ofN/R. In embodi(cid:173)
`ments, the ratio N/R is between 8 and 28. In further embodi(cid:173)
`ments, the ratio N/R is between 12 and 20 and more specifi(cid:173)
`cally may be between 15 and 16. The table below shows 40
`additional examples of solidity and the ratio N/R for different
`numbers of propulsor blades 74.
`
`TABLE
`
`Number of Blades. Solidity and Ratio N/R
`
`Number of Blades (N)
`
`Solidity
`
`Ratio N/R
`
`16
`14
`12
`10
`16
`14
`12
`10
`16
`14
`12
`10
`16
`14
`12
`10
`16
`14
`12
`10
`16
`14
`
`1.1
`1.1
`1.1
`1.1
`1.02
`1.02
`1.02
`1.02
`0.89
`0.89
`0.89
`0.89
`0.76
`0.76
`0.76
`0.76
`0.63
`0.63
`0.63
`0.63
`0.6
`0.6
`
`14.55
`12.73
`10.91
`9.09
`15.69
`13.73
`11.76
`9.80
`17.98
`15.73
`13.48
`11.24
`21.05
`18.42
`15.79
`13.16
`25.40
`22.22
`19.05
`15.87
`26.67
`23.33
`
`50
`
`55
`
`60
`
`65
`
`What is claimed is:
`1. A gas turbine engine comprising:
`a spool;
`a turbine coupled to drive the spool;
`a propulsor coupled to be driven by said turbine through
`said spool; and
`a gear assembly coupled between said propulsor and said
`spool such that rotation of said spool drives said propul(cid:173)
`sor at a different speed than said spool,
`wherein said propulsor includes a hub and a row of pro-
`pulsor blades that extend from said hub, and said row
`includes a number (N) of said propulsor blades that is no
`more than 16, and the propulsor is located at an inlet of
`a bypass flow passage having a pressure ratio that is
`between 1.1 and 1.35 with regard to an inlet pressure and
`an outlet pressure of said bypass flow passage;
`wherein each of said propulsor blades extends radially
`between a root and a tip and in a chord direction between
`a leading edge and a trailing edge at the tip to define a
`chord dimension (CD), said row of propulsor blades
`defining a circumferential pitch (CP) with regard to said
`tips, wherein said row of propulsor blades has a solidity
`value (R) defined as CD/CP that is between 0.6 and 0.9,
`and a ratio ofN/R is between 8 and 16 or between 18 and
`28.
`2. The gas turbine engine as recited in claim 1, wherein said
`pressure ratio is between 1.2 and 1.3.
`3. The gas turbine engine as recited in claim 1, wherein the
`propulsor is located at an inlet of a core flow passage and a
`bypass flow passage that define a design bypass ratio of
`
`GE-1001.005
`
`

`
`US 9,121,412 B2
`
`5
`approximately 18 with regard to flow through said core flow
`passage and said bypass flow passage.
`4. The gas turbine engine as recited in claim 1, wherein the
`design pressure ratio is between 1.1 and 1.2.
`5. The gas turbine engine as recited in claim 1, further
`comprising a low pressure compressor section and a high
`pressure compressor section, and said turbine includes a low
`pressure turbine section and a high pressure turbine section,
`said low pressure compressor section and said low pressure
`turbine section are each coupled to be driven though said 10
`spool, and said high pressure compressor section and said
`high pressure turbine section are each coupled to be driven
`through another spool.
`6. The gas turbine engineasrecitedinclaim 1, wherein said 15
`ratio ofN/R is between 8 and 10.
`7. The gas turbine engine as recited in claim 1, wherein said
`number (N) of said propulsor blades is 10 to 16.
`8. The gas turbine engine as recited in claim 1, wherein said
`number (N) of said propulsor blades is even.
`
`6
`9. A propulsor of a gas turbine engine, the propulsor com(cid:173)
`prising:
`a rotor including a row of propulsor blades extending radi(cid:173)
`ally outwardly from a hub, each of said propulsor blades
`extending radially between a root and a tip and in a chord
`direction between a leading edge and a trailing edge at
`the tip to define a chord dimension (CD) at the tip of each
`propulsor blade, said row of propulsor blades defining a
`circumferential pitch (CP) with regard to the tips,
`wherein said row of propulsor blades has a solidity value
`(R) defined as CD/CP wherein the solidity value is
`between 0.6 and 0.9, and said row includes a number (N)
`of said propulsor blades that is no greater than 16 such
`that a ratio ofN/R is between 8 and 16 or between 18 and
`28.
`10. The propulsor as recited in claim 9, wherein said ratio
`ofN/R is between 12 and 16 or between 18 and 20.
`11. The propulsor as recited in claim 9, wherein said ratio
`ofN/R is between 15 and 16.
`* * * * *
`
`GE-1001.006