GEARED ARCHITECTURE FOR HIGH SPEED
AND SMALL VOLUME FAN DRIVE TURBINE
BACKGROUND
The present disclosure relates to a gas turbine engine, and more particularly to
a flexibie support structure for a geared architecture therefor.
Epicyclic gearboxes with planetary or star gear trains may be used in gas
turbine engines for their compact designs and efficient high gear reduction
capabilities. Planetary and star gear trains generally include three gear train elements:
10 a central sun gear, an outer ring gear with interi~agl ear teeth, and a plurality of planet
gears supported by a planet carrier between and in meshed engagement wit11 both the
sun gear and the ring gear. The gear train elements share a common longitudinal
central axis, about which at least two rotate. An advantage of epicyclic gear trains is
that a rotary input can be connected to any one of the three elements. One of the
I5 other two elements is then held stationary with respect to the other two to permit the
third to serve as an output.
In gas turbine engine applications, where a speed reduction transmission is
required, the central sun gear generally receives rotary input from the power plant,
the outer ring gear is generally held stationary and the planet gear carrier rotates in
20 the same direction as the sun gear to provide torque output at a reduced rotational
speed. In star gear trains, the planet carrier is held stationary and the output shaft is
driven by the ring gear in a direction opposite that of the sun gear.
During flight, light weight structural cases deflect with aero and maneuver
loads causing significant a~nounts of transverse deflection commonlg known as
25 backbone bending of the engine. This deflection may cause the individual sun or
planet gear's axis of rotation to lose parallelism with the central axis. This deflection
inay result in some misalignment at gear train journal bearings and at the gear teeth
mesh, which may lead to efficiency losses froin the misalignment and potential
reduced life from increases in the concentrated stresses.
3 0 Further, with the geared architecture as set forth above, the torque and speed
of the input into the gear is quite high.
SUMMARY
In a featured embodiment, a gas turbine engine has a fan shaft driving a fan, a
frame supporting the fan shaft, and a plurality of gears to drive the fan shaft. A
flexible support at least partially supports the plurality of gears. The flexible support
5 has a lesser stiffness than the frame. A first turbine section provides a drive input
into the plurality of gears. A second turbine section is also included. The first
turbine section has a first exit area at a first exit point and rotates at a first speed. The
second turbine section has a second exit area at a second exit point and rotates at a
second speed, which is faster than the first speed. A first performance quantity is
10 defined as the product of the first speed squared and the first area. A second
performance quantity is defined as the product of the second speed squared and the
second area. A ratio of the first performance quantity to the second perforlnance
quantity is between about 0.5 and about 1.5.
In another embodiment according to the previous embodiment, the ratio is
15 above or equal to about 0.8.
In another embodiment according to any of the previous embodiments, the
first turbine section has at least three stages.
In another embodiment according to any of the previous embodiments, the
first turbine section has up to six stages.
2 0 In another embodiment according to any of the previous embodiments, the
second turbine section has two or fewer stages.
In another embodiment according to any of the previous embodiments, a
pressure ratio across the first turbine section is greater than about 5: 1.
In another embodiment according to any of the previous embodiments, a ratio
25 of a thrust provided by the engine, to a volume of a turbine section including both the
high pressure turbine and the low pressure turbine is greater than or equal to about
1.5 and less than or equal to about 5.5 1bflinch2.
In another embodiment according to any of the previous embodiments, the
frame includes a frame lateral stiffness and a frame transverse stiffness. The flexible
30 support includes a flexible support transverse stiffness and a flexible support lateral
stiffness. The flexible support lateral stiffness is less than the frame lateral stifftiess
and the flexible support transverse stiffness is less than the frame transverse stiffness.
In another embodiment according to any of the previous embodiments, a
flexible coupling connects at least one of the plurality of gears to be driven by the
first turbine section.
In another embodiment according to any of the previous enlbodiments, the
5 flexible coupling has a flexible coupling lateral stiffness and a flexible coupling
transverse stiffness. The flexible coupling lateral stiffness is less than the frame
lateral stiffness. The flexible coupling transverse stiffness is less than the frame
transverse stiffness.
In another embodiment according to ally of the previous embodiments, the
10 plurality of gears include a gear mesh that defines a gear mesh lateral stiffness and a
gear mesh transverse stiffness. The gear mesh lateral stiffness is greater than the
flexible support lateral stiffness. The gear mesh transverse stiffness is greater than
the flexible support transverse stiffness.
In another featured embodiment, a gas turbine engine has a fan shaft driving a
15 fan, a frame which supports the fan shaft, and a plurality of gears which drives the
fan shaft. A flexible support which at least partially supports the plurality of gears
has a lesser stiffness than the frame. A high pressure turbine and a low pressure
turbine are included, the low pressure turbine being configured to drive one of the
plurality of gears. A ratio of a thrust provided by the engine, to a volume of a turbine
20 section including both the high pressure turbine and the low pressure turbine, is are
greater than or equal to about 1.5 and less than or equal to about 5.5 1bflinc1i2.
In another embodiment accordiilg to the previous embodiment, the ratio is
greater than or equal to about 2.0.
In another ernbodiilleilt according to any of the previous embodiments, the
25 ratio is greater than or equal to about 4.0.
In another embodiment according to any of the previous embodiments, the
thrust is sea level take-off, flat-rated static thrust.
In another embodiment according to any of the previous embodiments, the
frame includes a frame lateral stiffness and a fiame transverse stiffness. The flexible
30 support includes a flexible support transverse stiffness and a flexiblc support lateral
stiffness. The flexible support lateral stiffness is less than the frame lateral stiffness
and the flexible support transverse stiffiless is less than the frame transverse stiffness.
In another e~nbodiment according to any of the previous en~bodiments, a
flexible coupling connects at least one of the plurality of gears to be driven by the
first turbine section.
In another embodiment according to any of the previous embodiments, the
5 flexible coupling has a flexible coupling lateral stiffness and a flexible coupling
transverse stiffhess. The flexible coupling lateral stiffness is less than the frame
lateral stiffness, and the flexible coupling transverse stiffness is less than the frame
transverse stiffness.
In another enlbodilnent according to any of the previous embodiments, the
10 plurality of gears include a gear mesh that defines a gear mesh lateral stiffness and a
gear mesh transverse stiffness. The gear mesh lateral stiffi~essi s greater than the
flexible support lateral stiffness. The gear mesh transverse stiffness is greater than
the flexible support transverse stiffness.
In another featured embodiment, a gas turbine engine has a fan shaft and a
15 frame which supports the fan shaft. The fraine defines at least one of a frame lateral
stiffness and a frame transverse stiffness. A gear system drives the fan shaft. A
flexible support at least partially supports the gear system. The flexible support
defines at least one of a flexible support lateral stiffness with respect to the frame
lateral stiffness and a flexible suppoi-t transverse stiffness with respect to the frame
20 transverse stiffness. An input coupling to the gear system defines at least one of an
input coupling lateral stiffness with respect to the frame lateral stiffness and an input
coupling transverse stiffness with respect to the frame transverse stiffness.
BRIEF DESCRIPTION OF THE DRAWINGS
Various features will become apparent to those skilled in the art fro111 the
following detailed description of the disclosed non-limiting embodiment. The
drawings that accompany the detailed description can be briefly described as follows:
Figure 1A is a schematic cross-section of a gas turbine engine;
Figure 1 B shows a feature of the Figure 1 A engine.
Figure 1 C shows another feature.
Figure 1D shows yet another feature.
Figure 2 is an enlarged cross-section of a section of the gas t~~rbineeng ine
which illustrates a fail drive gear systein (FDGS);
Figure 3 is a schematic view of a flex mount arrangement for one 11011-
limiting embodiment of the FDGS;
Figure 4 is a scl~ematicv iew of a flex inount asrangenlent for another nonlimiting
embodiment of the FDGS;
5 Figure 5 is a schematic view of a flex mount arrangement for another 11011-
limiting embodiment of a star system FDGS; and
Figure 6 is a schematic view of a flex mount arrangement for another nonlimiting
embodiment of a planetary system FDGS.
Figure 7 is a schematic view of a flex mount arrangement for another lion-
10 limiting embodiment of a star system FDGS; and
Figure 8 is a schematic view of a flex mount arrangement for anotl~er nonlimiting
embodiment of a planetary system FDGS.
DETAILED DESCRIPTION
15 Figure 1 schematically illustrates a gas turbine engine 20. The gas turbinc
engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a
fan section 22, a compressor section 24, a combustor section 26 and a turbine section
28. Alternative engines might include an augmentor section (not sliown) among
other systems or features. The fan section 22 drives air along a bypass flow path B in
20 a bypass duct defined within a nacelle 15, while the compressor section 24 drives air
along a core flow path C for compression and communication into the combustor
section 26 then expansion through the turbine section 28. Although depicted as a
two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, ~t
should be understood that the concepts described herein are not limited to use with
25 two-spool turbofans as the teachings may be applied to other types of turbine engines
including three-spool architectures.
The exemplary engine 20 generally includes a low speed spool 30 and a 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. It sl1ould be
30 understood that various bearing systems 38 at various locations may alternatively or
additionally be provided, and the location of bearing systems 38 may be varied as
appropriate to the application.
The low speed spool 30 generally includes an inner shaft 40 that
interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46.
The inner shaft 40 is connected to the fan 42 through a speed change mechanism,
which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to
5 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 compressor 52 and
high pressure turbine 54. A combustor 56 is arranged in exen~plary gas turbine 20
between the high pressure compressor 52 and the high pressure turbine 54. A midturbine
frame 57 of the engine static structure 36 is arranged generally between the
10 high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57
further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and
the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine
central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow is compressed by the low pressure compressor 44 then the
15 high pressure compressor 52, mixed and burned with fuel in the combustor 56, then
expanded over the high pressure turbine 54 and low pressure turbine 46. The midturbine
frame 57 includes airfoils 59 which are in the core airflow path C. The
turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed
spool 32 in response to the expansion. It will be appreciated that each of the
20 positions of the fan section 22, colnpressor section 24, combustor section 26, turbine
section 28, and fan drive gear system 48 may be varied. For example, gear system 48
may be located aft of combustor section 26 or even aft of turbine section 28, and fan
section 22 may be positioned forward or aft of the location of gear system 48.
Thc engine 20 in one example is a high-bypass geared aircraft engine. In a
25 further example, the engine 20 bypass ratio is greater than about six (6), with a11
example embodiment being greater than about ten (lo), the geared arcliitecture 48 is
an epicyclic gear train, such as a planetary gear system or other gear systcm, with a
gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a
pressure ratio that is greater than about five. In one disclosed embodiment, the
30 engine 20 bypass ratio is greater than about ten (10:1), the fan diaii~eter is
significantly larger than that of the low pressure compressor 44, and the low pressure
turbine 46 has a pressure ratio that is greater than about five 5:l. Low pressure
turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine
46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an
-7-
exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a
planetary gear system or other gear system, with a gear reduction ratio of greater than
about 2.3:l. It should be understood, however, that the above parameters are only
exemplary of one embodiment of a geared architecture engine and that the present
5 invention is applicable to other gas turbine engines including direct drive turbofans.
A significant amount of thrust is provided by the bypass flow B due to the
high bypass ratio. The fan section 22 of the engine 20 is designed for a particular
flight condition -- typically cruise at about 0.8 Mach and about 35,000 feet. The
flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel
10 consumption - also known as "bucket cruise Thrust Specific Fuel Consumption
('TSFC')" - is the industry standard parameter of Ibm of fuel being burned divided
by lbf of thrust the engine produces at that minimum point. "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
15 non-limiting embodiment is less than about 1.45. "Low corrected fan tip speed" is
the actual fan tip speed in fi/sec divided by an industry standard temperature
correction of [(Tram OR) / (518.7 OR)]' '. The "Low corrected fan tip speed" as
disclosed herein according to one non-limiting embodiment is less than about 1 150 St
/ second.
2 0 The core airflow is con~pressedb y the low pressure compressor 44 then the
high pressure compressor 52, mixed and burned with fuel in the combustor 56, then
expanded over the high pressure turbine 54 and low pressure turbine 46. The
turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed
spool 32 in response to the expansion of the airflow passing therethrough.
2 5 The amount of thrust that can be produced by a particular turbine section
compared to how compact the turbine section is, is referred to as the power density,
or the force density, of the turbine section, and is derived by the flat-rated Sea Level
Take-Off (SLTO) thrust divided by the volume of the entire turbine section. The
example volume is determined fi-om an inlet of the high pressure turbine 54 to an exit
30 of the low pressure turbine 46. In order to increasc the power density of the turbine
section 28, each of the low pressure and high pressure turbines 46, 54 is made n~ore
compact. That is, the high pressure turbine 54 and the low pressure turbine 46 are
made with a sllorter axial length, and the spacing between each of the turbines 46, 54
is decreased, thereby decreasing the voluil~eo f the turbine section 28.
-8-
The power density in the disclosed gas turbine engine 20 including the gear
driven fan section 22 is greater than those provided in prior art gas turbine engine
including a gear driven fan. Eight disclosed exemplary engines, which incorporate
turbine sections and fan sections driven througl~ a reduction gear system and
In some embodiments, the power density is greater than or equal to about I .5
5 architectures as set forth in this application, are described in Table I as follows:
TABLE 1
lbf /in3. In further embodiments, the power density is greater than or equal to about
Engine
1
2
3
4
5
6
7
8
10 2.0 lbf /in3. In further embodiments, the power density is greater than or equal to
about 3.0 lbf/in3. In further embodiments, the power density is greater than or equal
Thrust SLTO
(lbf)
17,000
23,300
29,500
3 3,000
96,500
96,500
96,500
37,098
to about 4.0 lbf /in3. In further embodiments, the power density is less than or equal
to about 5.5 lbf /in3,
Turbine section volume
from the Inlet
3,859
5,330
6,745
6,745
3 1,086
62,172
46,629
6,745
Engines made with the disclosed gear driven fan architecture, and including
Thrust/turbine section
volume (lbf /in3)
4.4
4.37
4.37
4.84
3.1
1.55
2.07
5.50
15 turbine sections as set forth in this application, provide very high efficiency
operation, and increased fuel efficiency.
Referring to Figure lB, with continued reference to Figure lA, relative
rotations between components of example disclosed engine architecture 100 are
schematically shown. In the example engine architecture 100, the fan 42 is
20 connectcd, through the gearbox 48, to the low spool 30 to which the low pressure
colnpressor 44 and the low pressure turbine 46 are connected. The high pressure
compressor 52 and the high pressure turbine 54 are connected to a CoInmoIl shaft
forming the high spool 32. The high spool 32 rotates opposite the direction of
rotation of the fan 42 (illustrated in Figure IB as the "+" direction.) The low spool 30
rotates in the same dircction as the fan 42 (illustrated in Figure IB as the "-"
direction.) The high pressure turbine 54 and the low pressure turbine 46, along with
the mid-turbine frame 57 together forms the turbine section 28 of the gas turbine
5 engine 20. Other relative rotation directions between the two spools and the fan
come within the scope of this disclosure.
One disclosed example speed change device 48 has a gear reduction ratio
exceeding 2.3:1, meaning that the low pressure turbine 46 turns at least 2.3 times
faster than the fan 42. An example disclosed speed change device is an epicyclical
10 gearbox of a planet type, where the input is to the center "sun" gear 260. Planet
gears 262 (only one shown) around the sun gear 260 rotate and are spaced apart by a
carrier 264 that rotates in a direction common to the sun gear 260. A ring gear 266,
which is non-rotatably fixed to the engine static casing 36 (shown in Figure I ) ,
contains the entire gear assembly. The fan 42 is attached to and driven by the carrier
15 264 such that the direction of rotation of the fan 42 is the same as the direction of
rotation of the carrier 264 that, in turn, is the same as the direction of rotation oi'the
input sun gear 260. Accordingly, the low pressure compressor 44 and the low
pressure turbine 46 counter-rotate relative to the high pressure compressor 52 and the
high pressure turbine 54.
20 Counter rotating the low pressure compressor 44 and the low pressure turbine
46 relative to the high pressure compressor 52 and the high pressure turbine 54
provides certain efficient aerodynamic conditions in the turbine section 28 as the
generated high speed exhaust gas flow moves from the high pressure turbine 54 to
the low pressure turbine 46. Moreover, the mid-turbine frame 57 contributes to the
25 overall compactness of the turbine section 28. Further, the airfoil 59 of the midturbine
frame 57 surrounds internal bearing support structures and oil tubes that are
cooled. The airfoil 59 also directs flow around the internal bcaring support structures
and oil tubes for streamlining the high speed exhaust gas flow. Additionally, the
airfoil 59 directs flow exiting the high pressure turbine 54 to a proper angle desired to
30 promote increased efficiency of the low pressure turbine 46.
Flow exiting the high pressure turbine 54 has a significant component of
tangential swirl. The flow direction exiting the high pressure turbine 54 is set almost
ideally for the blades in a first stage of the low pressure turbine 46 for a wide range
of engine power settings. Thus, the aerodynamic turning function of the mid turbine
-1 0-
frame 57 can be efficiently achieved without dramatic additional align~ilent of
airflow exiting the high pressure turbine 54.
Referring to Figure 1C, the example turbine section 28 volume is
schematically shown and includes first, second and third stages 46A, 46B and 46C.
5 Each of the stages 46A, 46B and 46C includes a corresponding plurality of blades
212 and vanes 214. The example turbine section further includes an example airturning
vane 220 between the low and high turbines 54, 46 that has a modest cainber
to provide a small degree of redirection and achieve a desired flow angle relative to
blades 212 of the first stage 46a of the low pressure turbine 46. The disclosed vane
10 220 could not efficiently perforin the desired airflow function if the low and high
pressure turbines 54, 46 rotated in a comlnon direction.
The example mid-turbine frame 57 includes multiple air turning vanes 220 in
a row that direct air flow exiting the high pressure turbine 54 and ensure that air is
flowing in the proper direction and with the proper amount of swirl. Because the
15 disclosed turbine section 28 is more compact than previously utilized turbine
sections, air has less distance to travel between exiting the mid-turbine fralne 57 and
entering the low pressure turbine 46. The smaller axial travel distance results in a
decrease in the amount of swirl lost by the airflow during the transition from the midturbine
frame 57 to the low pressure turbine 46, and allows the vanes 220 of the mid-
20 turbine frame 57 to function as inlet guide vanes of the low pressure turbine 46. The
mid-turbine frame 57 also includes a strut 221 providing structural suppost to both
the mid-turbine frame 57 and to the engine housing. In one example, the mid-turbine
frame 57 is much more compact by encasing the strut 221 within the vane 220.
thereby decreasing the length of the mid-turbine frame 57.
2 5 At a given fan tip speed and thrust level provided by a given fan size, thc
inclusion of'the speed change device 48 (shown in Figures 1A and 1B) provides a
gear reduction ratio, and thus the speed of the low pressurc turbine 46 and low
pressure compressor 44 components may be increased. More speciiically, for a
given fan diameter and fan tip speed, increases in gear ratios provide for a faster
30 turning turbine that, in turn, provides for an increasingly compact turbine and
increased thrust to volume ratios of the turbine section 28. By increasing the gear
reduction ratio, the speed at which the low pressure compressor 44 and the low
pressure turbine 46 turn, relative to the speed of the fan 42, is increased.
Increases in rotational speeds of the gas turbine engine 20 compone~its
increases overall efficiency, thereby providing for reductions in the diameter and the
number of stages of the low pressure compressor 44 and the low pressure turbine 46
that would otherwise be required to maintain desired flow characteristics of the air
5 flowing through the core flow path C. The axial length of each of the low pressure
compressor 44 and the low pressure turbine 46 can therefore be further reduced due
to efficiencies gained from increased speed provided by an increased gear ratio.
Moreover, the reduction in the diameter and the stage count of the turbine section 28
increases the compactness and provides for an overall decrease in required axial
10 length of the example gas turbine engine 20.
In order to further improve the thrust density of the gas turbine engine 20, the
example turbine section 28 (including the high pressure turbine 54, the mid-turbine
frame 57, and the low pressure turbine 46) is made more compact than traditional
turbine engine designs, thereby decreasing the length of the turbine section 28 and
15 the overall length of the gas turbine engine 20.
In order to make the example low pressure turbine 46 compact, make the
diameter of the low pressure turbine 46 more compatible with the high pressure
turbine 54, and thereby make the air-turning vane 220 of the mid-turbine frame 57
practical, stronger inaterials in the initial stages of the low pressure turbine 46 niay be
20 required. The speeds and centrifugal pull generated at the compact diameter of the
low pressure turbine 46 pose a challenge to materials used in prior art low pressure
turbines.
Examples of materials and processes within the contemplation of this
disclosure for the air-turning vane 220, the low pressure turbine blades 212, and tlie
25 vanes 214 include materials with directionally solidified grains to provided added
strength in a span-wise direction. An example method for creating a vane 220, 214
or turbine blade 212 having directionally solidified grains can be found in U.S.
Applications No. 131290667, and U.S. Patent Nos. 7338259 and 7871247, each of
which is incorporated by reference. A fusther, engine embodiment utilizes a cast,
30 hollow blade 212 or vane 214 with cooling air introduced at the leading edge of tlie
bladehane and a trailing edge discharge of the cooling air. Another embodin~ent
uses an internally cooled blade 212 or vane 214 wit11 film cooling holes. An
additional engine embodiment utilizes an aluininum lithiuln material for construction
of a portion of the low pressure turbine 46. The example low pressure turbine 46
may also be constructed utilizing at a powdered metal disc or rotor.
Additionally, one or more rows of turbine blades 212 of the low pressure
turbine 46 can be constructed using a single crystal blade material. Single crystal
5 constructions oxidize at higher temperatures as compared to non-single crystal
constructions and thus can withstand higher temperature airflow. Higher temperature
capability of the turbine blades 212 provide for a more efficient low pressure turbine
46 that may be further reduced in size.
While the illustrated low pressure turbine 46 includes three turbine stages
10 46a, 46b, and 46c, the low pressure turbine 46 can be modified to include up to six
turbine stages. Increasing the number of low pressure turbine stages 46a, 46b, 46c at
constant thrust slightly reduces the thrust density of the turbine section 28 but also
increases power available to drive the low pressure compressor and the fan section
22.
15 Further, the example turbine blades may be internally cooled to allow the
material to retain a desired strength at higher temperatures and thereby perforn~ as
desired in view of the increased centrifugal force generated by the conipact
configuration while also withstanding the higher temperatures created by adding low
pressure compressor 44 stages and increasing fan tip diameter.
20 Each of the disclosed embodiments enables tlie low pressure turbine 46 to be
more compact and efficient, while also inlproving radial alignment to the high
pressure turbine 54. Improved radial alignment between the low and high pressure
turbines 54, 46 increases efficiencies that can offset any increases in manufacturing
costs incurred by including the air turning vane 220 of the mid-turbine frame 57.
25 In light of the foregoing embodiments, the overall size of the turbine section
28 has been greatly reduced, thereby enhancing the engine's power density. Further,
as a result of the improvemelit in power density, the engine's ovcrall propulsive
efficiency has been improved.
An exit area 400 is shown, in Figure ID and Figure 1 A, at the exit location
30 for the high pressure turbine section 54. An exit area for tlie low pressure turbine
section is defined at exit 401 for the low pressure turbine section. As shown in
Figure ID, the turbine engine 20 may be counter-rotating. This means that the low
pressure turbine section 46 and low pressure coinpressor section 44 rotate in one
direction, while the high pressure spool 32, including high pressure turbine section
- 13-
54 and high pressure compressor section 52 rotate in a11 opposed direction. The gear
reduction 48, which may be, for example, an epicyclic transmission (e.g., with a sun,
ring, and star gears), is selected such that the fan 42 rotates in the same direction as
the high spool 32. With this arrangenzcnt, and with the other structure as set forth
5 above, including the various quantities and operational ranges, a very high speed can
be provided to the low pressure spool. Low pressure turbine section and high
pressure turbine section operation are often evaluated looking at a performance
quantity which is the exit area for the turbine section inultiplied by its respective
speed squared. This performance quantity ("PQ") is defined as:
10 Equalion 1: PQlt,, = (All,( x v,,~')
Equation 2: PQllPt = (All1,t x VI~,,?)
where Alpt is the area of the low pressure turbine section at the exit thereof
(e.g., at 401), where VIP, is the speed of the low pressure turbine section, where All,,,
is the area of the high pressure turbine section at the exit thereof (e.g., at 400), and
15 where Vllpt is the speed of the low pressure turbine section.
Thus, a ratio of the performance quantity for the low pressure turbine section
compared to the performance quantify for the high pressure turbine section is:
Equation 3: (AiPt x v~ptZ)/(~llxp t~ 11prZ)= PQltP/P Q11~1t
In one turbine embodiment made according to the above design, the areas of
20 the low and high pressure turbine sections are 557.9 in2 and 90.67 in2, respectively.
Further, the speeds of the low and high pressure turbine sections are 101 79 rpn~a nd
24346 rpm, respectively. Thus, using Equations 1 and 2 above, the perforlnance
quantities for the low and high pressure turbine sections are:
Equation 1: PQltp = (Alpt x vll,:) = (557.9 in2)(10179 r p ~ n )=~
25 57805 157673.9 in2 rpm2
Equation 2: PQilpt = (Allpt x Vh,,?) = (90.67 in2)(24346 rp111)~ -=
53742622009.72 in2 rpn12
and using Equation 3 above, the ratio for the low pressure turbine section to
the high pressure turbine section is:
3 0 Ratio = PQltpi PQllpi = 57805157673.9 in2 rpm2 153742622009.72 in2 rpm2 =
1.075
In another embodiment, the ratio was about 0.5 and in another e~nbodiment
the ratio was about 1.5. With PQltp, PQIlpt ratios in the 0.5 to 1.5 range, a very
efficient overall gas turbine engine is achieved. More narrowly, PQItplP QllPrta tios of
-14-
above or equal to about 0.8 are more efficient. Even more narrowly, PQl,,,/ PQI,~,,
ratios above or equal to 1.0 are even more efficient. As a result of these PQltp/ PQlIl,,
ratios, in particular, the turbine section can be made much smaller than in tlie prior
art, both in diameter and axial length. In addition, the efficiency of the overall engine
5 is greatly increased.
The low pressure compressor section is also itnproved with this arrangement,
and behaves more like a high pressure compressor section than a traditional low
pressure compressor section. It is more efficient than the prior art, and can provide
more work in fewer stages. The low pressure compressor section may be made
10 smaller in radius and shorter in length while contributing more toward achieving the
overall pressure ratio design target of the engine.
A worker of ordinary skill in the art, being apprised of the disclosure above,
would recognize that high torque and high speed will be presented by the low speed
spool 30 into the gear architecture 48. Thus, a flexible mount arralige~nentb ecomes
15 important.
With reference to Figure 2, the geared architecture 48 generally includes a fan
drive gear system (FDGS) 60 driven by the low speed spool 30 (illustrated
schematically) through an input coupling 62. The input coupling 62 both transfers
torque from the low speed spool 30 to the geared architecture 48 and facilitates the
20 segregation of vibrations and other transients therebetween. In the disclosed nonlimiting
embodiment, the FDGS 60 may include an epicyclic gear system which lvay
be, for example, a star system or a planet system.
The input coupling 62 may include an interface spline 64 joined, by a gear
spline 66, to a sun gear 68 of the FDGS 60. The sun gear 68 is in meshed
25 engagement with multiple planet gears 70, of which the illustrated planet gear 70 is
representative. Each planet gear 70 is rotatably mounted in a planet carrier 72 by a
respective planet journal bearing 75, Rotary motion of the sun gcar 68 urges each
planet gear 70 to rotate about a respective longitudinal axis P. The gears may be
generally as shown schematically in Figure 1 B.
3 0 Each planet gear 70 is also in meshed engagenlent wit11 rotating ring gear 74
that is mechanically connected to a fan shaft 76. Since the planet gears 70 mesh with
both the rotating ring gear 74 as well as tlie rotating sun gear 68, the planet gears 70
rotate about their own axes to drive tlie ring gear 74 to rotate about engine axis A.
The rotation of the ring gear 74 is conveyed to the fan 42 (Figure 1) througll the fan
-15-
shaft 76 to thereby drive the fan 42 at a lower speed than the low speed spool 30. It
should be understood that the described geared architecture 48 is but a single 11011-
limiting embodiment and that various other geared architectures will alternatively
benefit herefrom.
5 With reference to Figure 3, a flexible support 78 supports the planet carrier 72
to at least partially support the FDGS 60A with respect to the static structure 36 such
as a front center body which facilitates the segregation of vibrations and other
transients therebetween. It should be understood that various gas turbine engine case
structures may alternatively or additionally provide the static structure and flexible
10 silpport 78. It is to be understood that the term "lateral" as used herein refers to a
pel-pendicular direction with respect to the axis of rotation A and the term
"transverse" refers to a pivotal bending movement with respect to the axis of rotation
A so as to absorb deflections which may be otherwise applied to the FDGS 60. The
static structure 36 may further include a number 1 and 1.5 bearing support static
15 structure 82 which is commonly refelred to as a "K-frame" which supports the
number 1 and number 1.5 bearing systems 38A. 38B. Notably, the K-frame bearing
support defines a lateral stiffness (represented as Kframe in Figure 3) and a
transverse stiffness (represented as Kfralne REND in Figure 3) as the relerenced factors
in this non-limiting embodiment.
20 In this disclosed non-limiting emboditnent, the lateral stiffness (KFS; KIC) of
both the flexible support 78 and the input coupling 62 are each less than about 11%
of the lateral stiffness (Kframe). That is, the lateral stiffness of the entire FDGS 60 is
controlled by this lateral stiffness relationship. Alternatively, or in addition to this
relationship, the transverse stiffness of both the flexible support 78 and the input
25 coupling 62 are each less than about 11% of the transverse stiffness (Kframe F3FNI) ),
That is, the transverse stiffness of the entire FDGS 60 is controlled by this transverse
stiffness relationship.
With reference to Figure 4, another non-limiting embodinlent of a FDGS 60B
includes a flexible support 78' that supports a rotationally fixed ring gear 74'. The
30 fan shaft 76' is driven by the planet carrier 72' in the sche~naticallyi llustrated planet
system which otherwise generally follows the star system architecture of Figure 3.
With reference to Figure 5, the lateral stiffness relationship within a FDGS
60C itself (for a star system architecture) is schematically represented. The lateral
stiffness (KIC) of an input coupling 62, a lateral stiffness (KFS) of a flexible support
-1 6-
78, a lateral stiffness (KRG) of a ring gear 74 and a lateral stiffness (KJB) of a planet
journal bearing 75 are controlled with respect to a lateral stiffness (KGM) of a gear
niesh within the FDGS 60.
In the disclosed non-limiting embodiment, the stiffness (KGM) may be
5 defined by the gear mesh between the sun gear 68 and the multiple planet gears 70.
The lateral stiffness (KGM) within the FDGS 60 is the referenced factor and the
static structure 82' rigidly supports the fan shaft 76. That is, the fan shaft 76 is
supported upon bearing systems 38A, 38B which are essentially rigidly supported by
the static structure 82'. The lateral stiffness (KJB) may be mechanically defined by,
10 for example, the stiffness within the planet journal bearing 75 and the lateral stiffness
(KRG) of the ring gear 74 may be mechanically defined by, for example, the
geometry of the ring gear wings 74L, 74R (Figure 2).
In the disclosed non-limiting embodiment, the lateral stiffness (KKG) of the
ring gear 74 is less than about 12% of the lateral stiffness (KGM) of the gear mesh;
15 the lateral stiffness (KFS) of the flexible support 78 is less than about 8% of the
lateral stiffness (KGM) of the gear mesh; the lateral stiffness (KJB) of the planet
journal bearing 75 is less than or equal to the lateral stiffness (KGM) of the gear
mesh; and the lateral stiffness (KIC) of an input coupling 62 is less than about 5% of
the lateral stiffness (KGM) of the gear mesh.
2 0 With reference to Figure 6, another non-limiting embodiment of a lateral
stiffness relationship within a FDGS 60D itself are schematically illustrated for a
planetary gear system architecture, which otherwise generally follows the star system
architecture of Figure 5.
It should be understood that combinations of the above lateral stiffness
25 relationships may be utilized as well. The lateral stiffiiess of each of structural
components may be readily measured as colnpared to film stiffness and spline
stiffiless which may be relatively difficult to determine.
By flex mounting to accommodate misalignment of the shafts under design
loads, the FDGS design loads have been reduced by more than 17% which reduces
30 overall engine weight, The flex mount facilitates alignment to illcrease system life
and reliability. The lateral flexibility in the flexible support and input coupling allows
the FDGS to essentially 'float' with the fan shaft during maneuvers. This allows:
(a) the torque transmissions in the fan shaft, the input coupling and the flexible
support to remain constant during maneuvers; (b) maneuver induced lateral loads in
-1 7-
the fan shaft (which may otherwise potentially misalign gears and daiiiage teeth) to
be mainly reacted to through the number 1 and 1.5 bearing support K-frame; and (c)
both the flexible support and the input coupling to transmit small a~nountso f lateral
loads into the FDGS. The splines, gear tooth stiffness, journal bearings, and ring
5 gear ligaments are specifically designed to minimize gear tooth stress variations
during maneuvers. The other connections to the FDGS are flexible mounts (turbine
coupling, case flex mount). These mount spring rates have been determined from
analysis and proven in rig and flight testing to isolate the gears from engine
maneuver loads. In addition, the planet journal bearing spring rate may also be
10 controlled to support system flexibility.
Figure 7 is similar to Figure 5 but shows the transverse stifhess relationships
within the FDGS 60C (for a star system arcl~itecture). The transverse stiffness
(K1cBENDo)f tlie input coupling 62, a transverse stiffness (KFS'"') of the flexible
support 78, a transverse stiffness (KRGBENDo)f the ring gear 74 and a transverse
15 stiffness (K.lBBENDo)f the planet journal bearing 75 are controlled with respect to a
transverse stiffness ( K G M " ~o~f ~th)e gear mesh within the FDGS 60.
In the disclosed non-limiting embodiment, the stiffness (KGMR"") niay be
defined by the gear mesh between the sun gear 68 and tlie multiple planet gears 70.
The transverse stiffncss ( K G M ~w~ith~in~ th)e FDGS 60 is the referenced factor and
20 the static structurc 82' rigidly supports tlie fan shaft 76. That is, the fan shaft 76 is
supported upon bearing systems 38A, 38B which are essentially rigidly supported by
the static structure 82'. The transverse stiffness (KJB'"') may be niecliaiiically
defined by, for example, the stiffness within tlie planet journal bearing 75 and tlie
transverse stiffness (KRGBENDo)f the ring gear 74 may be mechanically defined by,
25 for example, the geometry of the ring gear wings 74L, 74R (Figure 2).
in the disclosed non-limiting embodiment, the transverse stiffness (KRG'"')
of the ring gear 74 is less than about 12% of the transverse stiffncss (KCM~"") ol'
the gear mesh; the transverse stiffness (KFSBmD) of the flexible support 78 is less
than about 8% of the transverse stiffiiess (KGM'"~) of the gear mesh; tlie transverse
30 stiffiiess ( K J B ~ o~f ~the~ p)la net journal bearing 75 is less than or equal to tlie
transverse stiffness ( K G M ' ~ ~o~f) the gear mesh; and the transverse stiffness
(KICBmD) of an input coupling 62 is less than about 5% of the transverse stiffness
(KGM""~") of the gear mesh.
Figure 8 is similar to Figure 6 but shows the transverse stiffness relationship
within the FDGS 60D for the planetary gear system architecture.
It should be understood that relative positional terms suc11 as "forward," "aft,"
"upper," "lower," "above," "below," and the like are with reference to the nor117al
5 operational attitude of the vehicle and should not be considered otherwise limiting.
It should be understood that like reference numerals identify corresponding or
similar elements throughout the several drawings. It should also be uilderstood that
although a particular component arrangement is disclosed in the illustrated
embodiment, other arrange~nelltsw ill benefit herefroin.
10 Although particular step sequences are shown, described, and claimed, it
should be understood that steps may be performed in any order, separated or
combined unless otherwise indicated and will still benefit from the present
disclosure.
The combined arrangement of the high power density and fan drive turbine
15 with the high AN^ performance quantity, all incorporated with the flexible mounting
structure, provide a very robust and efficient gas turbine engine.
The foregoing description is exemplary rather thail defined by the liinitatio~~s
within. Various non-limiting enlbodiments arc disclosed herein, however, one of
ordinary skill in the art would recogilize that various modifications and variations in
20 light of the above teachings will fall within the scope of the appended claims. It is
therefore to be understood that within the scope of the appended claims, the
disclosure may be practiced other than as specifically described. For that reason the
appended claims should be studied to determine true scope and content.
WE CLAIM:
1. A gas turbine engine comprising:
a fan shaft driving a fan;
a frame which supports said fan shaft;
a plurality of gears to drive said fan shaft;
a flexible support which at least partially supports said plurality of gears, said
flexible support having a lesser stiffness than said frame;
a first turbine section providing a drive input into said plurality of gears; and
a second turbine section,
wherein said first turbine section has a first exit area at a first exit point and
rotates at a first speed,
wherein said second turbine section has a second exit area at a second exit
point and rotates at a second speed, which is faster than the first speed,
15 wherein a first perfor~nance quantity is defined as the product of the first
speed squared and the first area,
wherein a second performance quantity is defined as the product of the
second speed squared and the second area, and
wherein a ratio of the first performance quantity to the second performance
20 quantity is between about 0.5 and about 1.5.
2. The turbine section as set forth in claim 1, wherein said ratio is above or
equal to about 0.8.
25 3. The turbltie section as set forth in claim 1, wherein said first turbine section
has at least 3 stages.
4. The turbine section as set forth in claim 1, wherein said first turbine section
has up to 6 stages.
3 0
5. The turbine section as set forth in claim 1, wherein said second turbine
section has 2 or fewer stages.
6. The turbine section as set forth in claim 1, wherein a pressure ratio across the
first turbine section is greater than about 5: 1.
7. The gas turbine engine as set forth in claiin 1, including a ratio of a thrust
5 provided by said engine, to a volume of a turbine section including both said high
pressure turbine and said low pressure turbine being greater than or equal to about
1.5 and less than or equal to about 5.5 1bf/incll2.
8. The gas turbine engine as set forth in claim 1, wherein said frame includes a
10 frame lateral stiffness and a frame transverse stiffness, and said flexible support
includes a flexible support transverse stiffi~essa nd a flexible suppoi-t lateral stiffiless,
and said flexible support lateral stiffness being less than said frame lateral stiffness
and said flexible support transverse stiffiless being less than said franle transverse
stiffness.
15
9. The gas turbine engine as set forth in claim 8, wherein a flexible coupling
connects at least one of said plurality of gears to be driven by said first turbine
section.
20 10. The gas turbine engine as set forth in claim 9, wherein said flexible coupling
has a flexible coupling lateral stiffness and a flexible coupling transverse stiffi~ess,
and said flexible coupling lateral stiffness being less than said frame lateral stiffi~ess,
and said flexible coupling transverse stiffness being less than said frame transverse
stiffness.
25
11. The gas turbine engine as set forth in claim 8, wherein said plurality of gears
include a gear mesh that defines a gear mesh lateral stiffness and a gcar mesh
transverse stiffness, said gear mesh lateral stiffness being greater than said flexible
support lateral stiffness and said gear mesh transverse stiffness being greater than
30 said flexible support transverse stiffi~ess.
12. A gas turbine engine comprising:
a fan shaft driving a fan;
a frame which supports said fan shaft;
a plurality of gears which drives the fan shaft;
5 a flexible support which at least partially supports said plurality of gears, said
flexible support having a lesser stiffiless than said frame;
a high pressure turbine;
a low pressure turbine, said low pressure turbine being configured to drive
one of said plurality of gears;
10 a ratio of a thrust provided by said engine, to a volume of a turbine section
including both said high pressure turbine and said low pressure turbine being greater
than or equal to about 1.5 and less than or equal to about 5.5 1bf/incli2.
13. The gas turbine engine as set forth in claim 12, wherein said ratio is greater
15 than or equal to about 2.0.
14. The gas turbine engine as set forth in claim 13, wherein said ratio is greater
than or equal to about 4.0.
20 15. The gas turbine engine as set forth in claiin 12, wherein said thrust is sea level
take-off, flat-rated static thrust.
16. The gas turbine engine as set forth in claim 12, wherein said frame includes a
frame lateral stiffness and a frame transverse stiffness, and said flexible support
25 includes a flexible support transverse stiffness and a flexible support lateral stiffi~ess.
and said flexible support lateral stiffness being less than said frame lateral stiffi~ess
and said flexible support transverse stiffness being less than said frame transverse
stiffness.
30 17. The gas turbine engine as set forth in claim 16, wherein a flexible coupling
connects at least one of said plurality of gears to be driven by said first turbine
section.
18. The gas turbine engine as set forth in claim 17, wherein said flexible coupling
has a flexible coupling lateral stiffness and a flexible coupling trailsverse stiffness,
and said flexible coupling lateral stiffness being less than said fraine lateral stiffness,
and said flexible coupling transverse stiffness being less than said frame transverse
5 stiffness.
19. The gas turbine engine as set forth in claim 16, wherein said plurality of gears
include a gear mesh that defines a gear mesh lateral stiffness and a gear mesl~
transverse stiffness, said gear mesh lateral stiffness being greater than said flexible
10 support lateral stiffness and said gear mesh transverse stiffness being greater tl1a11
said flexible support transverse stiffness.
20. A gas turbine engine coinprising:
a fan shaft;
15 a fsame which supports said fan shaft, said frame defines at least one of a
frame lateral stiffness and a frame transverse stiffness;
a gear system which drives said fan shaft;
a flexible support which at least partially supports said gear system, said
flexible support defines at least one of a flexible support lateral stiffness with respect
20 to said frame lateral stiffness and a flexible support transverse stiffness wit11 respect
to said frame transverse stiffness; and
an input coupling to said gear system, said input coupling defines at least one
of an input coupling lateral stiffness with respect to said frame lateral stiffness and an
input coupling transverse stiffness with respect to said frame transverse stiffness.
2 5
2 1. A gas turbine engine, substantially as herein described with reference to
accon~pallyirng drawings and examples.