(19)
(11)EP 2 904 239 B1

(12)EUROPEAN PATENT SPECIFICATION

(45)Mention of the grant of the patent:
06.05.2020 Bulletin 2020/19

(21)Application number: 13843578.9

(22)Date of filing:  12.02.2013
(51)Int. Cl.: 
F02C 7/20  (2006.01)
F02K 3/06  (2006.01)
B64D 27/12  (2006.01)
F02C 7/00  (2006.01)
(86)International application number:
PCT/US2013/025717
(87)International publication number:
WO 2014/055103 (10.04.2014 Gazette  2014/15)

(54)

ASSEMBLY COMPRISING A GEARED TURBOFAN, A PYLON AND A WING

ANORDNUNG UMFASSEND EINEN GETRIEBEFAN, EINE AUFHÄNGUNG UND EINE TRAGFLÄCHE

ASSEMBLAGE COMPRENANT UN MOTEUR À DOUBLE FLUX À ENGRENAGE, UN MÂT DE LIAISON ET UNE AILE


(84)Designated Contracting States:
AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR

(30)Priority: 02.10.2012 US 201261708927 P

(43)Date of publication of application:
12.08.2015 Bulletin 2015/33

(60)Divisional application:
20167383.7

(73)Proprietor: United Technologies Corporation
Farmington, CT 06032 (US)

(72)Inventor:
  • GUKEISEN, Robert L.
    Middletown, Connecticut 06457 (US)

(74)Representative: Dehns 
St. Bride's House 10 Salisbury Square
London EC4Y 8JD
London EC4Y 8JD (GB)


(56)References cited: : 
WO-A1-2014/055102
US-A1- 2004 128 978
US-A1- 2012 117 940
US-A1- 2012 124 964
US-A- 5 452 575
US-A1- 2008 191 088
US-A1- 2012 117 940
US-A1- 2012 233 982
  
  • "DESIGN STUDY OF AN AIR PUMP AND INTEGRAL LIFT ENGINE ALF-504 USING THE LYCOMING 502 CORE", , 31 July 1972 (1972-07-31), XP552733059, Retrieved from the Internet: URL:http://ntrs.nasa.gov/archive/nasa/casi .ntrs.nasa.gov/19730004744.pdf [retrieved on 2018-09-18]
  
Note: Within nine months from the publication of the mention of the grant of the European patent, any person may give notice to the European Patent Office of opposition to the European patent granted. Notice of opposition shall be filed in a written reasoned statement. It shall not be deemed to have been filed until the opposition fee has been paid. (Art. 99(1) European Patent Convention).


Description

BACKGROUND OF THE INVENTION



[0001] A minimum distance must be maintained between a bottom of a gas turbine engine and a runway, resulting in space limitations for wing mounted gas turbine engines. Larger landing gear can be employed to raise the aircraft, and therefore the gas turbine engine, relative to the runway. However, this can add weight to the aircraft. As fan section becomes larger, there are fewer options for mounting a gas turbine engine.

[0002] A prior art has turbine engine, having the features of the preamble of claim 1 is disclosed in US 2012/233982 A2. A prior art gas turbine engine is disclosed in US 2012/117940 A1.

SUMMARY OF THE INVENTION



[0003] According to a first aspect, the present invention provides an assembly as recited in claim 1.

[0004] In a further embodiment of the foregoing assembly, the geared architecture includes an epicyclic gearbox.

[0005] In a further embodiment of any of the foregoing assemblies, the pylon includes a forward portion and an aft portion. The fan is attached to the forward portion of the pylon. The turbine section includes the fan drive turbine which is attached to the aft portion of the pylon.

[0006] In a further embodiment of any of the foregoing assemblies, the turbine section is a low pressure turbine.

[0007] In a further embodiment of any of the foregoing assemblies, the turbine to fan diameter ratio is substantially 35% to substantially 40%.

[0008] In a further embodiment of any of the foregoing assemblies, the fan extends forward of the wing.

[0009] In a further embodiment of any of the foregoing assemblies a distance of substantially 11 inches (0.279 metres) is defined between the wing and an upper portion of the gas turbine engine

[0010] These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description.

BRIEF DESCRIPTION OF THE DRAWINGS



[0011] 

Figure 1 illustrates a schematic view of an embodiment of a gas turbine engine;

Figure 2 illustrates a side view of the gas turbine engine mounted to a pylon; and

Figure 3 illustrates a front view of the gas turbine engine mounted to the pylon;


DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT



[0012] Figure 1 schematically illustrates an example gas turbine engine 20, such as a geared turbofan engine, that includes a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section 22 includes a fan 42 and drives air along a bypass flow path B while the compressor section 24 draws air in along a core flow path C where air is compressed and communicated to the combustor section 26. In the combustor section 26, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section 28 where energy is extracted and utilized to drive the fan section 22 and the compressor section 24.

[0013] Although the disclosed non-limiting embodiment depicts a geared turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with geared turbofans as the teachings may be applied to other types of traditional turbine engines. For example, the gas turbine engine 20 can have a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive the fan 42 via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.

[0014] The example gas turbine engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about a central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.

[0015] The low speed spool 30 generally includes an inner shaft 40 that connects the fan 42 and a low pressure (or first) compressor section 44 to a low pressure (or first) turbine section 46. The inner shaft 40 drives the fan 42 through a speed change device, such as a geared architecture 48, to drive the fan 42 at a lower speed than the low speed spool 30. The high-speed spool 32 includes an outer shaft 50 that interconnects a high pressure (or second) compressor section 52 and a high pressure (or second) turbine section 54. The inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the central longitudinal axis A. A propulsor section includes the fan 42 and a portion of the geared architecture 48.

[0016] A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. In one example, the high pressure turbine 54 includes at least two stages to provide a double stage high pressure turbine 54. In another example, the high pressure turbine 54 includes only a single stage. As used herein, a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure" compressor or turbine. A generator section includes the compressors 44 and 52, the combustor 56, and the turbines 46 and 54, as well as a portion of the geared architecture 48.

[0017] The example low pressure turbine 46 has a pressure ratio that is greater than about 5. The pressure ratio of the example low pressure turbine 46 is measured prior to an inlet of the low pressure turbine 46 as related to the pressure measured at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.

[0018] A mid-turbine frame 58 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 58 further supports bearing systems 38 in the turbine section 28 as well as setting airflow entering the low pressure turbine 46.

[0019] The air in the core flow path C is compressed by the low pressure compressor 44 then by the high pressure compressor 52 mixed with fuel and ignited in the combustor 56 to produce high speed exhaust gases that are then expanded through the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 58 includes vanes 60, which are in the core flow path C and function as an inlet guide vane for the low pressure turbine 46. Utilizing the vane 60 of the mid-turbine frame 58 as the inlet guide vane for low pressure turbine 46 decreases the length of the low pressure turbine 46 without increasing the axial length of the mid-turbine frame 58. Reducing or eliminating the number of vanes in the low pressure turbine 46 shortens the axial length of the turbine section 28. Thus, the compactness of the gas turbine engine 20 is increased and a higher power density may be achieved.

[0020] The disclosed gas turbine engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine 20 includes a bypass ratio greater than substantially six (6), with an example embodiment being greater than substantially ten (10). The example geared architecture 48 is an epicyclic gearbox, an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than substantially 2.3.

[0021] In one disclosed embodiment, the gas turbine engine 20 includes a bypass ratio greater than substantially ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor 44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture 48 and that the present disclosure is applicable to other gas turbine engines.

[0022] A significant amount of thrust is provided by the air in the bypass flow path B due to the high bypass ratio. The fan section 22 of the gas turbine engine 20 is designed for a particular flight condition -- typically cruise at substantially 0.8 Mach and substantially 35,000 feet (10,668 metres). The flight condition of 0.8 Mach and 35,000 ft. (10,668 m), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point.

[0023] "Low fan pressure ratio" is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane ("FEGV") system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than substantially 1.50. In another non-limiting embodiment the low fan pressure ratio is less than substantially 1.45.

[0024] The example gas turbine engine includes the fan 42 that comprises in one non-limiting embodiment less than substantially 26 fan blades. In another non-limiting embodiment, the fan section 22 includes less than substantially 20 fan blades. Moreover, in one disclosed embodiment the low pressure turbine 46 includes no more than substantially 6 turbine rotors. In another non-limiting example embodiment the low pressure turbine 46 includes substantially 3 turbine rotors. A ratio between the number of fan blades and the number of low pressure turbine rotors is between substantially 3.3 and substantially 8.6. The example low pressure turbine 46 provides the driving power to rotate the fan section 22 and therefore the relationship between the number of turbine rotors 34 in the low pressure turbine 46 and the number of blades in the fan section 22 disclose an example gas turbine engine 20 with increased power transfer efficiency.

[0025] Figures 2 and 3 illustrate the gas turbine engine 20 mounted to a pylon 74, which is mounted to a wing 64. The efficiency of the turbine driven geared architecture 48, disclosed above, enables the use and fabrication of a smaller low pressure turbine 46, both in diameter and in the number or overall stages, as compared to a direct drive engine that rotates at a less efficient speed. This allows for alternate and more efficient mounting configurations of the gas turbine engine 20.

[0026] A minimum distance 70 or clearance should be maintained between a bottom 66 of the gas turbine engine 20 and a runway 62 (or ground) and should be taken into consideration when determining a mounting configuration for the gas turbine engine 20. When the engine 20 is mounted to the wing 64, a core engine section, including the low pressure turbine 46, can be mounted under the wing 64, and the fan section 22 can disposed forward of the wing 64.

[0027] A diameter 68 of the low pressure turbine 46 is much smaller than a diameter 76 of the fan 42. A ratio of the diameter 68 of the low pressure turbine 46 to the diameter 76 of the fan 42 is defined as the turbine to fan diameter ratio. In one example, the turbine to fan diameter ratio is substantially less than 45%. In one example, the turbine to fan diameter ratio is substantially 25% to 45%. In one example, the turbine to fan diameter ratio is substantially 35% to substantially 40%. As the low pressure turbine 46 has a small diameter 68, this allows the central longitudinal axis A to be located closer (or "close coupled", which provides for a cost and weight savings) to the wing 64 than a correspondingly capable direct drive engine. The gas turbine engine 20 can also be located more aft due to the reduced diameter 68 of the low pressure turbine 46.

[0028] As the low pressure turbine 46 has a reduced diameter 68, the pylon 74 can also have a height 72 that is greater, or maintained at a given size, than a pylon 74 which mounts a correspondingly direct drive engine, while still maintaining the desired minimum distance 70 relative to the runway 62. The increase in height 72 provides for additional structural support and strength.

[0029] The pylon 74 includes a forward portion 90 and an aft portion 92. The fan 42 is attached to the forward portion 90 of the pylon 74, and the low pressure turbine 46 is attached to the aft portion 92 of the pylon 74.

[0030] As further shown in Figure 3, the wing 64 has a sheared or gull wing configuration. The wing 64 includes an up angle portion 78 and a flat portion 80. A distance 82, or gutter height, of substantially 11 inches should be maintained between the wing 64 and an upper portion 88 of the gas turbine engine 20. The up angle portion 78 provides additional space for the gas turbine engine 20 to be mounted under the wing 64, and even allows the mounting of a larger gas turbine engine 20, while still maintaining a desired distance 82 and the minimum distance 70 or clearance. That is, a reduced distance 82 between the gas turbine engine 20 and the wing 64, reducing interference drag.

[0031] The gear reduction, or geared architecture 48, can be considered part of the turbofan architecture without departing from the scope of the disclosed embodiments.


Claims

1. An assembly comprising a gas turbine engine (20), a pylon (74) and a wing (64), the gas turbine engine being assembled to the pylon (74) mounted to the wing (64), the gas turbine engine (20) comprising:

a mounting structure for mounting the engine (20) to the pylon (74); and

a propulsor section including a fan (42) having a fan diameter (76); and a geared architecture driving the fan (42); and

a turbine section (28) including a fan drive turbine (46) having a turbine diameter (68), the fan drive turbine (46) driving the geared architecture,

characterised in that:
the ratio of the turbine diameter (68) to the fan diameter (72) is less than 45% and in that the wing (64) has a gull wing configuration which includes an up angle portion (78) and a flat portion (80).


 
2. The assembly as recited in claim 1 wherein the geared architecture includes an epicyclic gearbox.
 
3. The assembly as recited in claim 1 or 2, wherein the pylon (74) includes a forward portion (90) and an aft portion (92), the fan (42) is attached to the forward portion (90) of the pylon (74), and the turbine section (28) including the fan drive turbine (46) is attached to the aft portion (92) of the pylon (74).
 
4. The assembly as recited in any preceding claim, wherein the turbine section (28) is a low pressure turbine (46).
 
5. The assembly as recited in any preceding claim, wherein the turbine to fan diameter ratio is 35% to 40%.
 
6. The assembly as recited in any preceding claim, wherein the pylon (74) includes a forward portion (90) and an aft portion (92);
wherein the fan (42) is attached to the forward portion (90) of the pylon (74) and the turbine section (28) is attached to the aft portion (92) of the pylon (74);
and further comprising:

a compressor section (24); and

a combustor (56) in fluid communication with the compressor section (24).


 
7. The assembly of any preceding claim, wherein the fan (42) extends forward of the wing (64).
 
8. The assembly of any preceding claim, wherein a distance of 11 inches (0.279 metres) is defined between the wing (64) and an upper portion of the gas turbine engine (20).
 


Ansprüche

1. Anordnung, die ein Gasturbinentriebwerk (20), eine Aufhängung (74) und eine Tragfläche (64) umfasst, wobei das Gasturbinentriebwerk an der Aufhängung (74) angeordnet ist, die an die Tragfläche (64) montiert ist, wobei das Gasturbinentriebwerk (20) Folgendes umfasst:

eine Montagestruktur für das Montieren des Triebwerks (20) an die Aufhängung (74); und

einen Vortriebsbereich, der einen Fan (42) mit einem Fandurchmesser (76) beinhaltet; und eine Getriebearchitektur, die den Fan (42) antreibt; und

einen Turbinenabschnitt (28), der eine Fanantriebsturbine (46) beinhaltet, die einen Turbinendurchmesser (68) aufweist, wobei die Fanantriebsturbine (46) die Getriebearchitektur antreibt,

dadurch gekennzeichnet, dass:
das Verhältnis des Turbinendurchmessers (68) zu dem Fandurchmesser (72) weniger als 45 % beträgt und dass die Tragfläche (64) eine Knickflügelkonfiguration aufweist, die einen Aufwärtswinkelbereich (78) und einen flachen Bereich (80) beinhaltet.


 
2. Anordnung nach Anspruch 1, wobei die Getriebearchitektur ein Planetengetriebe beinhaltet.
 
3. Anordnung nach Anspruch 1 oder 2, wobei die Aufhängung (74) einen vorderen Bereich (90) und einen hinteren Bereich (92) beinhaltet, wobei der Fan (42) an den vorderen Bereich (90) der Aufhängung (74) angebracht ist und wobei der Turbinenabschnitt (28), der die Fanantriebsturbine (46) beinhaltet, an den hinteren Bereich (92) der Tragfläche (74) angebracht ist.
 
4. Anordnung nach einem der vorhergehenden Ansprüche, wobei der Turbinenabschnitt (28) eine Niederdruckturbine (46) ist.
 
5. Anordnung nach einem der vorhergehenden Ansprüche, wobei das Verhältnis des Turbinen- zu dem Fandurchmesser 35 % bis 40 % Prozent beträgt.
 
6. Anordnung nach einem der vorhergehenden Ansprüche, wobei die Aufhängung (74) einen vorderen Bereich (90) und einen hinteren Bereich (92) beinhaltet;
wobei der Fan (42) an dem vorderen Bereich (90) der Aufhängung (74) und der Turbinenabschnitt (28) an dem hinteren Bereich (92) der Aufhängung (74) angebracht ist;
und ferner Folgendes umfasst:

einen Verdichterbereich (24); und

eine Brennkammer (56) in Fluidverbindung mit dem Verdichterbereich (24).


 
7. Anordnung nach einem der vorhergehenden Ansprüche, wobei der Fan (42) sich nach vorn zu der Tragfläche (64) erstreckt.
 
8. Anordnung nach einem der vorhergehenden Ansprüche, wobei ein Abstand von 11 Inch (0,279 Meter) zwischen der Tragfläche (64) und einem oberen Bereich des Gasturbinentriebwerks (20) definiert ist.
 


Revendications

1. Assemblage comprenant un moteur à turbine à gaz (20), un mât de liaison (74) et une aile (64), le moteur à turbine à gaz étant assemblé au mât de liaison (74) monté sur l'aile (64), le moteur à turbine à gaz (20) comprenant :

une structure de montage pour monter le moteur (20) sur le mât de liaison (74) ; et

une section de propulseur comportant une soufflante (42) ayant un diamètre de soufflante (76) ; et une architecture à engrenage entraînant la soufflante (42) ; et

une section de turbine (28) comportant une turbine d'entraînement de soufflante (46) ayant un diamètre de turbine (68), la turbine d'entraînement de soufflante (46) entraînant l'architecture à engrenage,

caractérisé en ce que :
le rapport entre le diamètre de turbine (68) et le diamètre de soufflante (72) est inférieur à 45 % et en ce que l'aile (64) a une configuration d'aile de mouette qui comporte une partie d'angle supérieure (78) et une partie plate (80).


 
2. Assemblage selon la revendication 1, dans lequel l'architecture à engrenage comporte un engrenage épicycloïdal.
 
3. Assemblage selon la revendication 1 ou 2, dans lequel le mât de liaison (74) comporte une partie avant (90) et une partie arrière (92), la soufflante (42) est fixée à la partie avant (90) du mât de liaison (74), et la section de turbine (28) comportant la turbine d'entraînement de soufflante (46) est fixée à la partie arrière (92) du mât de liaison (74).
 
4. Assemblage selon une quelconque revendication précédente, dans lequel la section de turbine (28) est une turbine basse pression (46).
 
5. Assemblage selon une quelconque revendication précédente, dans lequel le rapport entre la turbine et le diamètre de soufflante est compris entre 35 % et 40 %.
 
6. Assemblage selon une quelconque revendication précédente, dans lequel le mât de liaison (74) comporte une partie avant (90) et une partie arrière (92) ;
dans lequel la soufflante (42) est fixée à la partie avant (90) du mât de liaison (74) et la section de turbine (28) est fixée à la partie arrière (92) du mât de liaison (74) ;
et comprenant en outre :

une section de compresseur (24) ; et

une chambre de combustion (56) en communication fluidique avec la section de compresseur (24).


 
7. Assemblage selon une quelconque revendication précédente, dans lequel la soufflante (42) s'étend à l'avant de l'aile (64).
 
8. Assemblage selon une quelconque revendication précédente, dans lequel une distance de 11 pouces (0,279 mètre) est définie entre l'aile (64) et une partie supérieure du moteur à turbine à gaz (20).
 




Drawing









REFERENCES CITED IN THE DESCRIPTION



This list of references cited by the applicant is for the reader's convenience only. It does not form part of the European patent document. Even though great care has been taken in compiling the references, errors or omissions cannot be excluded and the EPO disclaims all liability in this regard.

Patent documents cited in the description