BACKGROUND OF THE INVENTION
[0001] The present application relates to a gas turbine engine having an improved fuel consumption
based upon a combination of operational parameters.
[0002] Gas turbine engines are known, and typically include a fan which drives air into
a bypass duct, and also into a compressor section. The air is compressed in the compressor
section, and delivered into a combustor section where it is mixed with fuel and burned.
Products of this combustion pass downstream over turbine rotors, driving the turbine
rotors to rotate.
[0003] In the past, a low pressure turbine has rotated at a given speed, and driven a low
pressure compressor, and the fan at the same rate of speed. More recently, gear reductions
have been included such that the fan in a low pressure compressor can be driven at
different speeds.
[0004] A prior art gas turbine engine and method of operating such having the features of
the preamble to claims 1 and 5 is disclosed in
US 2011/0120078. Other prior art gas turbine engines and methods of operating such are disclosed
in
US 5259187.
SUMMARY OF THE INVENTION
[0005] From one aspect, the present invention provides a gas turbine engine in accordance
with claim 1.
[0006] In an embodiment of the above engine, the gear ratio is less than or equal to about
4.2.
[0007] In a further embodiment of the above engine or previous embodiment, the expansion
ratio is greater than or equal to about 5.7.
[0008] In a further embodiment of the above engine or previous embodiments, the fan has
an outer diameter that is greater than an outer diameter of the low pressure turbine
section.
[0009] From another aspect, the present invention provides a method of operating a gas turbine
engine in accordance with claim 5.
[0010] In a further embodiment of the above method, the gear reduction is less than or equal
to 4.2.
[0011] In a further embodiment of the above method or previous embodiments, the fan has
an outer diameter that is greater than an outer diameter of the low pressure turbine
section.
BRIEF DESCRIPTION OF THE DRAWINGS
[0012] The various features and advantages of this invention will become apparent to those
skilled in the art from the following detailed description of the currently preferred
embodiment. The drawings that accompany the detailed description can be briefly described
as follows:
Figure 1A is a general schematic partial fragmentary view of an exemplary gas turbine
engine embodiment for use with the present invention;
Figure 1B is a perspective side partial fragmentary view of a FEGV system which provides
a fan variable area nozzle;
Figure 2A is a sectional view of a single FEGV airfoil outside the scope of the present
invention;
Figure 2B is a sectional view of the FEGV illustrated in Figure 2A shown in a first
position;
Figure 2C is a sectional view of the FEGV illustrated in Figure 2A shown in a rotated
position;
Figure 3A is a sectional view of an embodiment of a single FEGV airfoil within the
scope of the present invention;
Figures 3B is a sectional view of the FEGV illustrated in Figure 3A shown in a first
position;
Figure 3C is a sectional view of the FEGV illustrated in Figure 3A shown in a rotated
position;
Figure 4A is a sectional view of a single FEGV slatted airfoil in an arrangement not
claimed;
Figures 4B is a sectional view of the FEGV illustrated in Figure 4A shown in a first
position; and
Figure 4C is a sectional view of the FEGV illustrated in Figure 4A shown in a rotated
position.
DETAILED DESCRIPTION
[0013] Figure 1 illustrates a general partial fragmentary schematic view of a gas turbofan
engine 10 suspended from an engine pylon P within an engine nacelle assembly N as
is typical of an aircraft designed for subsonic operation.
[0014] The turbofan engine 10 includes a core section within a core nacelle 12 that houses
a low spool 14 and high spool 24. The low spool 14 includes a low pressure compressor
16 and low pressure turbine 18. The low spool 14 drives a fan section 20 directly
or through a gear train 22. The high spool 24 includes a high pressure compressor
26 and high pressure turbine 28. A combustor 30 is arranged between the high pressure
compressor 26 and high pressure turbine 28. The low and high spools 14, 24 rotate
about an engine axis of rotation A.
[0015] The engine 10 in the disclosed embodiment is a high-bypass geared turbofan aircraft
engine in which the engine 10 bypass ratio is greater than ten (10), the turbofan
diameter is significantly larger than that of the low pressure compressor 16, and
the low pressure turbine 18 has a pressure, or expansion, ratio greater than five
(5). The gear train 22 may be an epicycle gear train such as a planetary gear system
or other gear system with a gear reduction ratio of greater than 2.5. It should be
understood, however, that the above parameters are exemplary of only one geared turbofan
engine and that the present invention is likewise applicable to other gas turbine
engines including direct drive turbofans.
[0016] Airflow enters a fan nacelle 34, which may at least partially surrounds the core
nacelle 12. The fan section 20 communicates airflow into the core nacelle 12 for compression
by the low pressure compressor 16 and the high pressure compressor 26. Core airflow
compressed by the low pressure compressor 16 and the high pressure compressor 26 is
mixed with the fuel in the combustor 30 then expanded over the high pressure turbine
28 and low pressure turbine 18. The turbines 28, 18 are coupled for rotation with
respective spools 24, 14 to rotationally drive the compressors 26, 16 and, through
the gear train 22, the fan section 20 in response to the expansion. A core engine
exhaust E exits the core nacelle 12 through a core nozzle 43 defined between the core
nacelle 12 and a tail cone 32.
[0017] A bypass flow path 40 is defined between the core nacelle 12 and the fan nacelle
34. The engine 10 generates a high bypass flow arrangement with a bypass ratio in
which approximately 80 percent of the airflow entering the fan nacelle 34 becomes
bypass flow B. The bypass flow B communicates through the generally annular bypass
flow path 40 and may be discharged from the engine 10 through a fan variable area
nozzle (FVAN) 42 which defines a variable fan nozzle exit area 44 between the fan
nacelle 34 and the core nacelle 12 at an aft segment 34S of the fan nacelle 34 downstream
of the fan section 20.
[0018] Referring to Figure 1B, the core nacelle 12 is generally supported upon a core engine
case structure 46. A fan case structure 48 is defined about the core engine case structure
46 to support the fan nacelle 34. The core engine case structure 46 is secured to
the fan case 48 through a multiple of circumferentially spaced radially extending
fan exit guide vanes (FEGV) 50. The fan case structure 48, the core engine case structure
46, and the multiple of circumferentially spaced radially extending fan exit guide
vanes 50 which extend therebetween is typically a complete unit often referred to
as an intermediate case. It should be understood that the fan exit guide vanes 50
may be of various forms. The intermediate case structure in the disclosed embodiment
includes a variable geometry fan exit guide vane (FEGV) system 36.
[0019] Thrust is a function of density, velocity, and area. One or more of these parameters
can be manipulated to vary the amount and direction of thrust provided by the bypass
flow B. A significant amount of thrust is provided by the bypass flow B due to the
high bypass ratio. The fan section 20 of the engine 10 is nominally designed for a
particular flight condition -- typically cruise at 0.8M and 35,000 feet (10,668 m).
[0020] As the fan section 20 is efficiently designed at a particular fixed stagger angle
for an efficient cruise condition, the FEGV system 36 and/or the FVAN 42 is operated
to adjust fan bypass air flow such that the angle of attack or incidence of the fan
blades is maintained close to the design incidence for efficient engine operation
at other flight conditions, such as landing and takeoff. The FEGV system 36 and/or
the FVAN 42 may be adjusted to selectively adjust the pressure ratio of the bypass
flow B in response to a controller C. For example, increased mass flow during windmill
or engine-out, and spoiling thrust at landing. Furthermore, the FEGV system 36 will
facilitate and in some instances replace the FVAN 42, such as, for example, variable
flow area is utilized to manage and optimize the fan operating lines which provides
operability margin and allows the fan to be operated near peak efficiency which enables
a low fan pressure-ratio and low fan tip speed design; and the variable area reduces
noise by improving fan blade aerodynamics by varying blade incidence. The FEGV system
36 thereby
provides optimized engine operation over a range of flight conditions with respect
to performance and other operational parameters such as noise levels.
[0021] Referring to Figure 2A, in arrangements outside the scope of the present invention,
but useful for understanding the invention, each fan exit guide vane 50 includes a
respective airfoil portion 52 defined by an outer airfoil wall surface 54 between
the leading edge 56 and a trailing edge 58. The outer airfoil wall 54 typically has
a generally concave shaped portion forming a pressure side and a generally convex
shaped portion forming a suction side. It should be understood that respective airfoil
portion 52 defined by the outer airfoil wall surface 54 may be generally equivalent
or separately tailored to optimize flow characteristics.
[0022] Each fan exit guide vane 50 is mounted about a vane longitudinal axis of rotation
60. The vane axis of rotation 60 is typically transverse to the engine axis A, or
at an angle to engine axis A. It should be understood that various support struts
61 or other such members may be located through the airfoil portion 52 to provide
fixed support structure between the core engine case structure 46 and the fan case
structure 48. The axis of rotation 60 may be located about the geometric center of
gravity (CG) of the airfoil cross section. An actuator system 62 (illustrated schematically;
Figure 1A), for example only, a unison ring operates to rotate each fan exit guide
vane 50 to selectively vary the fan nozzle throat area (Figure 2B). The unison ring
may be located, for example, in the intermediate case structure such as within either
or both of the core engine case structure 46 or the fan case 48 (Figure 1A).
[0023] In operation, the FEGV system 36 communicates with the controller C to rotate the
fan exit guide vanes 50 and effectively vary the fan nozzle exit area 44. Other control
systems including an engine controller or an aircraft flight control system may also
be usable with the present invention. Rotation of the fan exit guide vanes 50 between
a nominal position and a rotated position selectively changes the fan bypass flow
path 40. That is, both the throat area (Figure 2B) and the projected area (Figure
2C) are varied through adjustment of the fan exit guide vanes 50. By adjusting the
fan exit guide vanes 50 (Figure 2C), bypass flow B is increased for particular flight
conditions such as during an engine-out condition. Since less bypass flow will spill
around the outside of the fan nacelle 34, the maximum diameter of the fan nacelle
required to avoid flow separation may be decreased. This will thereby decrease fan
nacelle drag during normal cruise conditions and reduce weight of the nacelle assembly.
Conversely, by closing the FEGV system 36 to decrease flow area relative to a given
bypass flow, engine thrust is significantly spoiled to thereby minimize or eliminate
thrust reverser requirements and further decrease weight and packaging requirements.
It should be understood that other arrangements as well as essentially infinite intermediate
positions are likewise usable with the present invention.
[0024] By adjusting the FEGV system 36 in which all the fan exit guide vanes 50 are moved
simultaneously, engine thrust and fuel economy are maximized during each flight regime.
By separately adjusting only particular fan exit guide vanes 50 to provide an asymmetrical
fan bypass flow path 40, engine bypass flow may be selectively vectored to provide,
for example only, trim balance, thrust controlled maneuvering, enhanced ground operations
and short field performance.
[0025] Referring to Figure 3A, in accordance with the invention, the FEGV system 36' includes
a multiple of fan exit guide vane 50' which each includes a fixed airfoil portion
66F and pivoting airfoil portion 66P which pivots relative to the fixed airfoil portion
66F. The pivoting airfoil portion 66P includes a leading edge flap which is actuatable
by an actuator system 62' as described above to vary both the throat area (Figure
3B) and the projected area (Figure 3C).
[0026] Referring to Figure 4A, in an arrangement not claimed the FEGV system 36" includes
a multiple of slotted fan exit guide vane 50" which each includes a fixed airfoil
portion 68F and pivoting and sliding airfoil portion 68P which pivots and slides relative
to the fixed airfoil portion 68F to create a slot 70 vary both the throat area (Figure
4B) and the projected area (Figure 4C) as generally described above. This slatted
vane method not only increases the flow area but also provides the additional benefit
that when there is a negative incidence on the fan exit guide vane 50" allows air
flow from the high-pressure, convex side of the fan exit guide vane 50" to the lower-pressure,
concave side of the fan exit guide vane 50" which delays flow separation.
[0027] The use of the gear reduction 22 allows control of a number of operational features
in combination to achieve improved fuel efficiency. The expansion ratio (or pressure
ratio) across the low pressure turbine, which is the pressure entering the low pressure
turbine section divided by the pressure leaving the low pressure turbine section is
greater than or equal to about 5.0.
[0028] In another embodiment, it is greater than or equal to about 5.7. In this same combination,
the bypass ratio is greater than 10.0. The gear reduction ratio is greater than 2.5.
In an embodiment, the gear reduction ratio is less than or equal to about 4.2.
[0029] This combination provides a low pressure turbine section that can be very compact,
and sized for very high aerodynamic efficiency with a small number of stages (3 to
5, in accordance with the present invention). Further, the maximum diameter of these
stages can be minimized to improve installation clearance under the wings of an aircraft.
[0030] The foregoing description is exemplary rather than defined by the limitations within.
Many modifications and variations of the present invention are possible in light of
the above teachings. The preferred embodiments of this invention have been disclosed,
however, one of ordinary skill in the art would recognize that certain modifications
would come within the scope of this invention. It is, therefore, to be understood
that within the scope of the appended claims, the invention may be practiced otherwise
than as specifically described. For that reason the following claims should be studied
to determine the true scope and content of this invention.
1. A gas turbine engine (10) comprising:
a core section defined about an axis (A);
a fan section (20) mounted at least partially around said core section to define a
fan bypass flow path (40), wherein a bypass ratio for the gas turbine engine (10)
which compares the air being delivered by the fan section (20) into a bypass duct
to the amount of air delivered into the core section is greater than 10, an expansion
ratio across a low pressure turbine section (18) is greater than 5, and the low pressure
turbine section (18) drives the fan section (20) through a gear reduction (22), with
the gear reduction (22) having a ratio greater than 2.5; and
a multiple of fan exit guide vanes (50') in communication with said fan bypass flow
path (40),
characterised in that:
said multiple of fan exit guide vanes (50') are rotatable about an axis of rotation
(A) to vary an effective fan nozzle exit area for said fan bypass flow path (40);
said multiple of fan exit guide vanes (50') are independently rotatable;
said multiple of fan exit guide vanes (50') are simultaneously rotatable;
said multiple of fan exit guide vanes (50') are mounted within an intermediate engine
case structure (46,48,50); and
each of said multiple of fan exit guide vanes (50') includes a pivotable portion (66P)
rotatable about said axis of rotation (A) relative a fixed portion (66F), said pivotable
portion (66P) including a leading edge flap;
wherein the low pressure turbine section (18) has 3 to 5 stages.
2. The gas turbine engine as set forth in claim 1, wherein said gear ratio is less than
or equal to about 4.2.
3. The gas turbine engine as set forth in claim 1 or 2, wherein said expansion ratio
is greater than or equal to about 5.7.
4. The gas turbine engine as set forth in any preceding claim, wherein said fan (20)
has an outer diameter that is greater than an outer diameter of the low pressure turbine
section (18).
5. A method of operating a gas turbine engine (10) including the steps of:
driving a fan (20) to deliver a first portion of air into a bypass duct, and a second
portion of air into a low pressure compressor (16), a bypass ratio of the first portion
to the second portion being greater than or equal to 8.0;
the first portion of air being delivered into the low pressure compressor (16), into
a high pressure compressor (26), and then into a combustion section (30), the air
being mixed with fuel and ignited, and products of the combustion passing downstream
over a high pressure turbine (28), and then a low pressure turbine (18), the low pressure
turbine section (18) being operated with an expansion ratio greater than or equal
to 5.0; and
said low pressure turbine section (18) being driven to rotate, and in turn rotating
said low pressure compressor (16), and rotating said fan (20) through a gear reduction
(22), said gear reduction (22) having a ratio of greater than or equal to 2.5, wherein
the gas turbine engine comprises a fan section (20) mounted at least partially around
a core section to define a fan bypass flow path (40),
characterised in that:
said fan section (20) comprises a multiple of fan exit guide vanes (50') rotatable
about an axis of rotation (A) to vary an effective fan nozzle exit area for said fan
bypass flow path (40);
said multiple of fan exit guide vanes (50') are independently rotatable;
said multiple of fan exit guide vanes (50') are simultaneously rotatable;
said multiple of fan exit guide vanes (50') are mounted within an intermediate engine
case structure (46,48,50); and
each of said multiple of fan exit guide vanes (50') includes a pivotable portion (66P)
rotatable about said axis of rotation (A) relative a fixed portion (66F), said pivotable
portion (66P) including a leading edge flap;
wherein the low pressure turbine section (18) has 3 to 5 stages.
6. The method as set forth in claim 5, wherein said gear reduction (22) is less than
or equal to 4.2.
7. The method as set forth in claim 5 or 6, wherein said fan (20) has an outer diameter
that is greater than an outer diameter of the low pressure turbine section (18).
1. Gasturbinentriebwerk (10), umfassend:
einen Kernabschnitt, der um eine Achse (A) definiert ist;
einen Gebläseabschnitt (20), der zumindest teilweise um den Kernabschnitt montiert
ist, um einen Gebläsenebenstromweg (40) zu definieren, wobei ein Nebenstromverhältnis
für das Gasturbinentriebwerk (10), das die Luft, die von dem Gebläseabschnitt (20)
in einen Nebenkanal geliefert wird, mit der Menge an Luft vergleicht, die in den Kernabschnitt
geliefert wird, vergleicht, größer als 10 ist, ein Expansionsverhältnis über einem
Niederdruckturbinenabschnitt (18) größer als 5 ist und der Niederdruckturbinenabschnitt
(18) den Gebläseabschnitt (20) über eine Räderuntersetzung (22) antreibt, wobei die
Räderuntersetzung (22) ein Verhältnis von über 2,5 aufweist; und
eine Vielzahl von Gebläseaustrittsleitschaufeln (50') in Kommunikation mit dem Gebläsenebenstromweg
(40),
dadurch gekennzeichnet, dass:
die Vielzahl von Gebläseaustrittsleitschaufeln (50') um eine Rotationsachse (A) drehbar
ist, um eine wirksame Gebläsedüsenaustrittsfläche für den Gebläsenebenstromweg (40)
zu variieren;
wobei die Vielzahl von Gebläseaustrittsleitschaufeln (50') unabhängig drehbar ist;
wobei die Vielzahl von Gebläseaustrittsleitschaufeln (50') gleichzeitig drehbar ist;
wobei die Vielzahl der Gebläseaustrittsleitschaufeln (50') innerhalb einer Zwischentriebwerksgehäusestruktur
(46, 48, 50) montiert ist; und
jede aus der Vielzahl der Gebläseaustrittsleitschaufeln (50') einen schwenkbaren Abschnitt
(66P) umfasst, der um die Rotationsachse (A) relativ zu einem feststehenden Abschnitt
(66F) drehbar ist, wobei der schwenkbare Abschnitt (66P) eine Vorderkantenklappe umfasst;
wobei der Niederdruckturbinenabschnitt (18) 3 bis 5 Stufen aufweist.
2. Gasturbinentriebwerk nach Anspruch 1, wobei das Übersetzungsverhältnis kleiner oder
gleich etwa 4,2 ist.
3. Gasturbinentriebwerk nach Anspruch 1 oder 2, wobei das Expansionsverhältnis größer
oder gleich etwa 5,7 ist.
4. Gasturbinentriebwerk nach einem der vorstehenden Ansprüche, wobei das Gebläse (20)
einen Außendurchmesser aufweist, der größer als ein Außendurchmesser des Niederdruckturbinenabschnitts
(18) ist.
5. Verfahren zum Betreiben eines Gasturbinentriebwerks (10), umfassend die folgenden
Schritte:
Antreiben eines Gebläses (20), um einen ersten Teil Luft in einen Nebenkanal zu liefern
und einen zweiten Teil Luft in einen Niederdruckverdichter (16) zu liefern, wobei
ein Nebenstromverhältnis des ersten Teils zum zweiten Teil größer oder gleich 8,0
ist;
wobei der erste Teil Luft in den Niederdruckverdichter (16), in einen Hochdruckverdichter
(26) und dann in einen Verbrennungsabschnitt (30) geliefert wird, die Luft mit Kraftstoff
gemischt und entzündet wird und die Produkte der Verbrennung stromabwärts über eine
Hochdruckturbine (28) und dann eine Niederdruckturbine (18) strömen, wobei der Niederdruckturbinenabschnitt
(18) mit einem Expansionsverhältnis von größer oder gleich 5,0 betrieben wird; und
wobei der Niederdruckturbinenabschnitt (18) betrieben wird, um sich zu drehen und
wiederum den Niederdruckverdichter (16) zu drehen und das Gebläse (20) über eine Räderuntersetzung
(22) zu drehen, wobei die Räderuntersetzung (22) ein Verhältnis von größer oder gleich
2,5 aufweist, wobei das Gasturbinentriebwerk einen Gebläseabschnitt (20) umfasst,
der zumindest teilweise um einen Kernabschnitt montiert ist, um einen Gebläsenebenstromweg
(40) zu definieren,
dadurch gekennzeichnet, dass:
der Gebläseabschnitt (20) eine Vielzahl von Gebläseaustrittsleitschaufeln (50') umfasst,
die um eine Rotationsachse (A) drehbar ist, um eine wirksame Gebläsedüsenaustrittsfläche
für den Gebläsenebenstromweg (40) zu variieren;
wobei die Vielzahl von Gebläseaustrittsleitschaufeln (50') unabhängig drehbar ist;
wobei die Vielzahl von Gebläseaustrittsleitschaufeln (50') gleichzeitig drehbar ist;
wobei die Vielzahl der Gebläseaustrittsleitschaufeln (50') innerhalb einer Zwischentriebwerksgehäusestruktur
(46, 48, 50) montiert ist; und
jede aus der Vielzahl der Gebläseaustrittsleitschaufeln (50') einen schwenkbaren Abschnitt
(66P) umfasst, der um die Rotationsachse (A) relativ zu einem feststehenden Abschnitt
(66F) drehbar ist, wobei der schwenkbare Abschnitt (66P) eine Vorderkantenklappe umfasst;
wobei der Niederdruckturbinenabschnitt (18) 3 bis 5 Stufen aufweist.
6. Verfahren nach Anspruch 5, wobei die Räderuntersetzung (22) kleiner oder gleich 4,2
ist.
7. Verfahren nach Anspruch 5 oder 6, wobei das Gebläse (20) einen Außendurchmesser aufweist,
der größer als ein Außendurchmesser des Niederdruckturbinenabschnitts (18) ist.
1. Turbine à gaz (10) comprenant :
une section de noyau définie autour d'un axe (A) ;
une section de ventilateur (20) montée au moins partiellement autour de ladite section
de noyau pour définir un trajet d'écoulement de dérivation de ventilateur (40), dans
laquelle un rapport de dérivation pour la turbine à gaz (10) qui compare l'air délivré
par la section de ventilateur (20) dans un conduit de dérivation à la quantité d'air
délivrée dans la section de noyau est supérieur à 10, un rapport de détente sur une
section de turbine à basse pression (18) est supérieur à 5, et la section de turbine
à basse pression (18) entraîne la section de ventilateur (20) par un réducteur à engrenages
(22), le réducteur à engrenages (22) ayant un rapport supérieur à 2,5 ; et
une pluralité d'aubes de guidage de sortie de ventilateur (50') en communication avec
ledit trajet d'écoulement de dérivation de ventilateur (40),
caractérisée en ce que :
ladite pluralité d'aubes de guidage de sortie de ventilateur (50') peut tourner autour
d'un axe de rotation (A) pour faire varier une zone de sortie de buse de ventilateur
efficace pour ledit trajet d'écoulement de dérivation de ventilateur (40) ;
ladite pluralité d'aubes de guidage de sortie de ventilateur (50') peut tourner indépendamment
;
ladite pluralité d'aubes de guidage de sortie de ventilateur (50') peut tourner simultanément
;
ladite pluralité d'aubes de guidage de sortie de ventilateur (50') est montée à l'intérieur
d'une structure de carter de moteur intermédiaire (46, 48, 50) ; et
chacune de ladite pluralité d'aubes de guidage de sortie de ventilateur (50') comprend
une partie pivotante (66P) pouvant tourner autour dudit axe de rotation (A) par rapport
à une partie fixe (66F), ladite partie pivotante (66P) comprenant un volet de bord
d'attaque ;
dans laquelle la section de turbine à basse pression (18) a 3 à 5 étages.
2. Turbine à gaz selon la revendication 1, dans laquelle ledit rapport d'engrenage est
inférieur ou égal à environ 4,2.
3. Turbine à gaz selon la revendication 1 ou 2, dans laquelle ledit rapport de détente
est supérieur ou égal à environ 5,7.
4. Turbine à gaz selon une quelconque revendication précédente, dans laquelle ledit ventilateur
(20) a un diamètre extérieur qui est supérieur à un diamètre extérieur de la section
de turbine à basse pression (18).
5. Procédé de fonctionnement d'une turbine à gaz (10) comprenant les étapes :
d'entraînement d'un ventilateur (20) pour délivrer une première partie d'air dans
un conduit de dérivation, et une seconde partie d'air dans un compresseur à basse
pression (16), un rapport de dérivation de la première partie à la seconde partie
étant supérieur ou égal à 8,0 ;
la première partie de l'air étant délivrée dans le compresseur à basse pression (16),
dans un compresseur à haute pression (26), puis dans une section de combustion (30),
l'air étant mélangé avec du carburant et enflammé, et les produits de la combustion
passant en aval sur une turbine à haute pression (28), puis une turbine à basse pression
(18), la section de turbine à basse pression (18) fonctionnant avec un rapport de
détente supérieur ou égal à 5,0 ; et
ladite section de turbine à basse pression (18) étant entraînée en rotation et à son
tour faisant tourner ledit compresseur à basse pression (16) et faisant tourner ledit
ventilateur (20) par l'intermédiaire d'un réducteur à engrenages (22), ledit réducteur
à engrenages (22) ayant un rapport supérieur ou égal à 2,5, dans lequel la turbine
à gaz comprend une section de ventilateur (20) montée au moins partiellement autour
d'une section de noyau pour définir un trajet d'écoulement de dérivation de ventilateur
(40),
caractérisé en ce que :
ladite section de ventilateur (20) comprend une pluralité d'aubes de guidage de sortie
de ventilateur (50') pouvant tourner autour d'un axe de rotation (A) pour faire varier
une zone de sortie de buse de ventilateur efficace pour ledit trajet d'écoulement
de dérivation de ventilateur (40) ;
ladite pluralité d'aubes de guidage de sortie de ventilateur (50') peut tourner indépendamment
;
ladite pluralité d'aubes de guidage de sortie de ventilateur (50') peut tourner simultanément
;
ladite pluralité d'aubes de guidage de sortie de ventilateur (50') est montée à l'intérieur
d'une structure de carter de moteur intermédiaire (46, 48, 50) ; et
chacune de ladite pluralité d'aubes de guidage de sortie de ventilateur (50') comprend
une partie pivotante (66P) pouvant tourner autour dudit axe de rotation (A) par rapport
à une partie fixe (66F), ladite partie pivotante (66P) comprenant un volet de bord
d'attaque ;
dans lequel la section de turbine à basse pression (18) a 3 à 5 étages.
6. Procédé selon la revendication 5, dans lequel ledit réducteur à engrenages (22) est
inférieure ou égale à 4,2.
7. Procédé selon la revendication 5 ou 6, dans lequel ledit ventilateur (20) a un diamètre
extérieur qui est supérieur à un diamètre extérieur de la section de turbine à basse
pression (18).