U.S. patent application number 11/496676 was filed with the patent office on 2010-07-22 for flade fan with different inner and outer airfoil stagger angles at a shroud therebetween.
This patent application is currently assigned to General Electric Company. Invention is credited to John Jared Decker, Aaron Michael Dziech, Peter Nicholas Szucs, Alan Glen Turner, Aspi Rustom Wadia.
Application Number | 20100180572 11/496676 |
Document ID | / |
Family ID | 38828693 |
Filed Date | 2010-07-22 |
United States Patent
Application |
20100180572 |
Kind Code |
A1 |
Wadia; Aspi Rustom ; et
al. |
July 22, 2010 |
FLADE FAN WITH DIFFERENT INNER AND OUTER AIRFOIL STAGGER ANGLES AT
A SHROUD THEREBETWEEN
Abstract
A FLADE fan assembly includes radially inner and outer airfoils
extending radially inwardly and outwardly respectively from an
annular shroud circumferentially disposed about a centerline. Inner
and outer chords extend between inner and outer leading and
trailing edges of inner and outer airfoil cross-sections of the
radially inner and outer airfoils respectively. Inner and outer
stagger angles between the inner and outer chords respectively at
the shroud and the centerline are different. The radially outer
airfoils may outnumber the radially inner airfoils and particularly
by a ratio in a range of 1.5:1 to about 4:1. Load paths or radii
may extend radially through the inner and outer airfoils and
through the rotating shroud between the inner and outer airfoils
and may pass near or through the inner and outer leading edges and
through the inner and outer trailing edges.
Inventors: |
Wadia; Aspi Rustom;
(Loveland, OH) ; Turner; Alan Glen; (Cincinnati,
OH) ; Dziech; Aaron Michael; (Cincinnati, OH)
; Szucs; Peter Nicholas; (7262 Adene Hills Ct West
Chester, OH) ; Decker; John Jared; (16607 Cottage
Rose Trail Cypress, TX) |
Correspondence
Address: |
STEVEN J. ROSEN
4729 CORNELL ROAD
CINCINNATI
OH
45241
US
|
Assignee: |
General Electric Company
|
Family ID: |
38828693 |
Appl. No.: |
11/496676 |
Filed: |
July 31, 2006 |
Current U.S.
Class: |
60/226.1 ;
60/39.162 |
Current CPC
Class: |
Y02T 50/60 20130101;
Y02T 50/673 20130101; F02K 3/072 20130101; Y02T 50/671 20130101;
F02K 3/077 20130101; F01D 5/022 20130101 |
Class at
Publication: |
60/226.1 ;
60/39.162 |
International
Class: |
F02K 3/072 20060101
F02K003/072; F02C 3/00 20060101 F02C003/00 |
Goverment Interests
[0001] The Government has rights in this invention pursuant to
Contract No. F33615-03-D-2352 awarded by the Department of Defense.
Claims
1. A FLADE counter-rotating fan aircraft gas turbine engine
comprising: axially spaced-apart forward and aft counter-rotatable
fans circumferentially disposed about a centerline, at least one
row of FLADE fan blades having radially outer airfoils disposed
radially outwardly of and drivingly connected to one of the forward
and aft counter-rotatable fans having radially inner airfoils, an
annular shroud disposed between the radially inner and outer
airfoils, the radially inner and outer airfoils extending radially
inwardly and outwardly respectively from the annular shroud, a core
engine and in flow receiving relationship with and located
downstream of the radially inner airfoils, and a bypass duct
surrounding the core engine and in flow receiving relationship with
and located downstream of the radially outer airfoils, inner and
outer chords extending between inner and outer leading and trailing
edges of inner and outer airfoil cross-sections of the radially
inner and outer airfoils respectively, inner and outer stagger
angles between the inner and outer chords respectively at the
annular shroud and the centerline, and the inner and outer stagger
angles being different.
2. An engine as claimed in claim 1 further comprising the radially
outer airfoils outnumbering the radially inner airfoils.
3. An engine as claimed in claim 2 further comprising the radially
outer airfoils outnumbering the radially inner airfoils by a ratio
in a range of 1.5:1 to about 4:1.
4. An engine as claimed in claim 2 further comprising the radially
outer airfoils outnumbering the radially inner airfoils by a ratio
of 2:1.
5. An engine as claimed in claim 2 further comprising the radially
outer airfoils outnumbering the radially inner airfoils by a ratio
of 1.5:1.
6. An engine as claimed in claim 2 further comprising radially
extending linear load paths extend radially along radii from the
centerline through the inner and outer airfoils and through the
annular shroud between the inner and outer airfoils.
7. An engine An assembly as claimed in claim 6 further comprising
the radially outer airfoils outnumbering the radially inner
airfoils by a ratio in a range of 1.5:1 to about 4:1.
8. An engine as claimed in claim 6 further comprising the radially
outer airfoils outnumbering the radially inner airfoils by a ratio
of 2:1.
9. An engine as claimed in claim 6 further comprising the radially
outer airfoils outnumbering the radially inner airfoils by a ratio
of 1.5:1.
10. An engine as claimed in claim 6 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, a first portion of the radially
extending linear load paths passing near or through the inner and
outer leading edges, and a second portion of the radially extending
linear load paths passing near or through the inner and outer
trailing edges.
11. An engine as claimed in claim 6 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, a first portion of the radially
extending linear load paths passing near or through the inner and
outer leading edges, a second portion of the radially extending
linear load paths passing near or through the inner and outer
trailing edges, and a third portion of the radially extending
linear load paths passing through inner and outer points between
the inner and outer trailing edges along the inner and outer chords
respectively.
12. An engine as claimed in claim 6 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, and multiple portions of the radially
extending linear load paths passing through inner and outer points
between the inner and outer trailing edges along the inner and
outer chords respectively.
13. An engine as claimed in claim 2 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, at least a first circumferential row
of radii extending radially outwardly from and normal to the
centerline, and each radius of the radii intersecting one of the
inner chords and one of the outer chords.
14. An engine as claimed in claim 13 further comprising: a second
circumferential row of the radii extending radially outwardly from
and normal to the centerline, the radii in the first
circumferential row passing near or through the inner and outer
leading edges, and the radii in the second circumferential row
passing near or through the inner and outer trailing edges.
15. An engine as claimed in claim 14 further comprising a third
circumferential row of the radii extending radially outwardly from
and normal to the centerline and the radii in the third
circumferential row passing between the inner and outer trailing
edges along the inner and outer chords respectively.
16. (canceled)
17. A FLADE counter-rotating fan aircraft gas turbine engine
comprising: axially spaced-apart forward and aft counter-rotatable
fans circumferentially disposed about a centerline, at least one
row of FLADE fan blades having radially outer airfoils disposed
radially outwardly of and drivingly connected to one of the forward
and aft counter-rotatable fans having radially inner airfoils, an
annular shroud disposed between the radially inner and outer
airfoils, the radially inner and outer airfoils extending radially
inwardly and outwardly respectively from the annular shroud, inner
and outer chords extending between inner and outer leading and
trailing edges of inner and outer airfoil cross-sections of the
radially inner and outer airfoils respectively, inner and outer
stagger angles between the inner and outer chords respectively at
the annular shroud and the centerline, the inner and outer stagger
angles being different, and the radially outer airfoils
outnumbering the radially inner airfoils.
18. An engine as claimed in claim 17 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio in a range of 1.5:1 to about 4:1.
19. An engine as claimed in claim 17 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio of 2:1.
20. An engine as claimed in claim 17 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio of 1.5:1.
21. An engine as claimed in claim 17 further comprising radially
extending linear load paths extend radially along radii from the
centerline through the inner and outer airfoils and through the
annular shroud between the inner and outer airfoils.
22. An engine as claimed in claim 21 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio in a range of 1.5:1 to about 4:1.
23. An engine as claimed in claim 21 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio of 2:1.
24. An engine as claimed in claim 21 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio of 1.5:1.
25. An engine as claimed in claim 21 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, a first portion of the radially
extending linear load paths passing near or through the inner and
outer leading edges, and a second portion of the radially extending
linear load paths passing near or through the inner and outer
trailing edges.
26. An engine as claimed in claim 21 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, a first portion of the radially
extending linear load paths passing near or through the inner and
outer leading edges, a second portion of the radially extending
linear load paths passing near or through the inner and outer
trailing edges, and a third portion of the radially extending
linear load paths passing through inner and outer points between
the inner and outer trailing edges along the inner and outer chords
respectively.
27. An engine as claimed in claim 21 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, and multiple portions of the radially
extending linear load paths passing through inner and outer points
between the inner and outer trailing edges along the inner and
outer chords respectively.
28. An engine as claimed in claim 17 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, at least a first circumferential row
of radii extending radially outwardly from and normal to the
centerline, and each radius of the radii intersecting one of the
inner chords and one of the outer chords.
29. An engine as claimed in claim 28 further comprising: a second
circumferential row of the radii extending radially outwardly from
and normal to the centerline, the radii in the first
circumferential row passing near or through the inner and outer
leading edges, and the radii in the second circumferential row
passing near or through the inner and outer trailing edges.
30. An engine as claimed in claim 29 further comprising a third
circumferential row of the radii extending radially outwardly from
and normal to the centerline and the radii in the third
circumferential row passing between the inner and outer trailing
edges along the inner and outer chords respectively.
31. (canceled)
32. A FLADE counter-rotating fan aircraft gas turbine engine
comprising: axially spaced-apart forward and aft counter-rotatable
fans circumferentially disposed about a centerline, at least one
row of FLADE fan blades having radially outer airfoils disposed
radially outwardly of and drivingly connected to the aft
counter-rotatable fan having radially inner airfoils, an annular
shroud disposed between the radially inner and outer airfoils, the
radially inner and outer airfoils extending radially inwardly and
outwardly respectively from the annular shroud, inner and outer
chords extending between inner and outer leading and trailing edges
of inner and outer airfoil cross-sections of the radially inner and
outer airfoils respectively, inner and outer stagger angles between
the inner and outer chords respectively at the annular shroud and
the centerline, and the inner and outer stagger angles being
different and the radially outer airfoils outnumbering the radially
inner airfoils.
33. An engine as claimed in claim 32 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio in a range of 1.5:1 to about 4:1.
34. An engine as claimed in claim 32 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio of 2:1.
35. An engine as claimed in claim 32 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio of 1.5:1.
36. An engine as claimed in claim 32 further comprising radially
extending linear load paths extend radially along radii from the
centerline through the inner and outer airfoils and through the
annular shroud between the inner and outer airfoils.
37. An engine as claimed in claim 36 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio in a range of 1.5:1 to about 4:1.
38. An engine as claimed in claim 36 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio of 2:1.
39. An engine as claimed in claim 36 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio of 1.5:1.
40. An engine as claimed in claim 36 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, a first portion of the radially
extending linear load paths passing near or through the inner and
outer leading edges, and a second portion of the radially extending
linear load paths passing near or through the inner and outer
trailing edges.
41. An engine as claimed in claim 36 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, a first portion of the radially
extending linear load paths passing near or through the inner and
outer leading edges, a second portion of the radially extending
linear load paths passing near or through the inner and outer
trailing edges, and a third portion of the radially extending
linear load paths passing through inner and outer points between
the inner and outer trailing edges along the inner and outer chords
respectively.
42. An engine as claimed in claim 36 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, and multiple portions of the radially
extending linear load paths passing through inner and outer points
between the inner and outer trailing edges along the inner and
outer chords respectively.
43. An engine as claimed in claim 32 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, at least a first circumferential row
of radii extending radially outwardly from and normal to the
centerline, and each radius of the radii intersecting one of the
inner chords and one of the outer chords.
44. An engine as claimed in claim 43 further comprising: a second
circumferential row of the radii extending radially outwardly from
and normal to the centerline, the radii in the first
circumferential row passing near or through the inner and outer
leading edges, and the radii in the second circumferential row
passing near or through the inner and outer trailing edges.
45. An engine as claimed in claim 44 further comprising a third
circumferential row of the radii extending radially outwardly from
and normal to the centerline and the radii in the third
circumferential row passing between the inner and outer trailing
edges along the inner and outer chords respectively.
46. An aircraft gas turbine engine comprising: a FLADE fan
circumferentially disposed about a centerline and having at least
one row of FLADE fan blades disposed in a FLADE duct, the FLADE fan
including radially inner and outer airfoils extending radially
inwardly and outwardly respectively from an annular shroud
circumferentially disposed about a centerline, inner and outer
chords extending between inner and outer leading and trailing edges
of inner and outer airfoil cross-sections of the radially inner and
outer airfoils respectively, inner and outer stagger angles between
the inner and outer chords respectively at the annular shroud and
the centerline, and the inner and outer stagger angles being
different, a core engine and in flow receiving relationship with
and located downstream of the radially inner airfoils, and a bypass
duct surrounding the core engine and in flow receiving relationship
with and located downstream of the radially outer airfoils.
47. An engine as claimed in claim 46 further comprising the
radially outer airfoils outnumbering the radially inner
airfoils.
48. An engine as claimed in claim 47 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio in a range of 1.5:1 to about 4:1.
49. An engine as claimed in claim 47 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio of 2:1.
50. An engine as claimed in claim 47 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio of 1.5:1.
51. An engine as claimed in claim 47 further comprising radially
extending linear load paths extend radially along radii from the
centerline through the inner and outer airfoils and through the
annular shroud between the inner and outer airfoils.
52. An engine as claimed in claim 51 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio in a range of 1.5:1 to about 4:1.
53. An engine as claimed in claim 51 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio of 2:1.
54. An engine as claimed in claim 51 further comprising the
radially outer airfoils outnumbering the radially inner airfoils by
a ratio of 1.5:1.
55. An engine as claimed in claim 51 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, a first portion of the radially
extending linear load paths passing near or through the inner and
outer leading edges, and a second portion of the radially extending
linear load paths passing near or through the inner and outer
trailing edges.
56. An engine as claimed in claim 51 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, a first portion of the radially
extending linear load paths passing near or through the inner and
outer leading edges, a second portion of the radially extending
linear load paths passing near or through the inner and outer
trailing edges, and a third portion of the radially extending
linear load paths passing through inner and outer points between
the inner and outer trailing edges along the inner and outer chords
respectively.
57. An engine as claimed in claim 51 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, and multiple portions of the radially
extending linear load paths passing through inner and outer points
between the inner and outer trailing edges along the inner and
outer chords respectively.
58. An engine as claimed in claim 47 further comprising: inner and
outer airfoil cross-sections of the inner and outer airfoils at the
annular shroud, inner and outer chords extending between inner and
outer leading and trailing edges of the inner and outer airfoil
cross-sections respectively, at least a first circumferential row
of radii extending radially outwardly from and normal to the
centerline, and each radius of the radii intersecting one of the
inner chords and one of the outer chords.
59. An engine as claimed in claim 58 further comprising: a second
circumferential row of the radii extending radially outwardly from
and normal to the centerline, the radii in the first
circumferential row passing near or through the inner and outer
leading edges, and the radii in the second circumferential row
passing near or through the inner and outer trailing edges.
60. An engine as claimed in claim 59 further comprising a third
circumferential row of the radii extending radially outwardly from
and normal to the centerline and the radii in the third
circumferential row passing between the inner and outer trailing
edges along the inner and outer chords respectively.
61. An engine as claimed in claim 46 further comprising variable
FLADE inlet guide vanes disposed axially forward and upstream of
the outer airfoils.
Description
BACKGROUND OF THE INVENTION
[0002] 1. Field of the Invention
[0003] This invention relates to FLADE aircraft gas turbine engines
and, more particularly, to such engines with counter-rotatable
fans.
[0004] 2. Description of Related Art
[0005] High performance variable cycle gas turbine engines are
being designed because of their unique ability to operate
efficiently at various thrust settings and flight speeds both
subsonic and supersonic. An important feature of the variable cycle
gas turbine engine which contributes to its high performance is its
capability of maintaining a substantially constant inlet airflow as
its thrust is varied. This feature leads to important performance
advantages under less than full power engine settings or maximum
thrust conditions, such as during subsonic cruise.
[0006] Counter-rotating fan gas turbine engines have also been
designed and tested because of their unique and inherent ability to
operate efficiently. Furthermore, counter-rotating fans powered by
counter-rotating turbines eliminate the need for stator vanes in
the fan section of the engine and at least one nozzle in the
turbine section of the engine. This significantly decreases the
weight of the engine. One issue regarding engine efficiency is the
desirability of equalizing fan rotor torque between the
counter-rotatable fans.
[0007] One particular type of variable cycle engine called a FLADE
engine (FLADE being an acronym for "fan on blade") is characterized
by an outer fan driven by a radially inner fan and discharging its
flade air into an outer fan duct which is generally co-annular with
and circumscribes an inner fan duct circumscribing the inner fan.
One such engine disclosed in U.S. Pat. No. 4,043,121, entitled "Two
Spool Variable Cycle Engine", by Thomas et al., provides a flade
fan and outer fan duct within which variable guide vanes control
the cycle variability by controlling the amount of air passing
through the flade outer fan duct.
[0008] Other high performance aircraft variable cycle gas turbine
FLADE engines capable of maintaining an essentially constant inlet
airflow over a relatively wide range of thrust at a given set of
subsonic flight ambient conditions such as altitude and flight Mach
No. in order to avoid spillage drag and to do so over a range of
flight conditions have been studied. This capability is
particularly needed for subsonic part power engine operating
conditions. Examples of these are disclosed in U.S. Pat. No.
5,404,713, entitled "Spillage Drag and Infrared Reducing Flade
Engine", U.S. Pat. No. 5,402,963, entitled "Acoustically Shielded
Exhaust System for High Thrust Jet Engines", U.S. Pat. No.
5,261,227, entitled "Variable Specific Thrust Turbofan Engine", and
European Patent No. EP0567277, entitled "Bypass Injector Valve For
Variable Cycle Aircraft Engines". Previously designed FLADE fans
had the inner and outer portions of a fladed blade close to
continuous in section properties at the transition region or shroud
separating the inner and outer portions or the inner blade and
outer FLADE fan blade. This in turn resulted in the same number of
the inner blade and the outer FLADE fan blades.
[0009] It is highly desirable to have a counter-rotating fan
aircraft gas turbine engine that can modulate bypass flow from a
fan section around a core engine to the bypass stream and to
effectively operate at high fan hub and bypass stream pressure
ratios to provide high specific thrust at takeoff and climb power
settings and to operate at low bypass stream pressure ratios to
provide good specific fuel consumption during reduced power cruise
operation. It is also desirable to provide counter-rotating fan
engines to eliminate the stator vanes in the fan section of the
engine, minimize the number of nozzles or vanes in the turbine, and
equalize fan rotor torque between the counter-rotatable fans. It is
also desirable to be able to design the inner fan blades and the
outer FLADE fan blades for maximum efficiency.
SUMMARY OF THE INVENTION
[0010] A FLADE fan assembly includes radially inner and outer
airfoils extending radially inwardly and outwardly respectively
from a rotatable annular shroud circumferentially disposed about a
centerline. Inner and outer chords extend between inner and outer
leading and trailing edges of inner and outer airfoil
cross-sections of the radially inner and outer airfoils
respectively. Inner and outer stagger angles between the inner and
outer chords respectively at the shroud and the centerline are
different.
[0011] The radially outer airfoils may outnumber the radially inner
airfoils and particularly by a ratio in a range of 1.5:1 to about
4:1. The FLADE fan assembly having the different inner and outer
stagger angles is particularly useful in a FLADE counter-rotating
fan aircraft gas turbine engine. The FLADE counter-rotating fan
aircraft gas turbine engine includes axially spaced-apart upstream
and downstream or forward and aft counter-rotatable fans
circumferentially disposed about a centerline, at least one row of
the FLADE fan blades having the radially outer airfoils disposed
radially outwardly of and drivingly connected to one of the forward
and aft counter-rotatable fans having radially inner airfoils. A
more particular embodiment of the FLADE counter-rotating fan
aircraft gas turbine engine includes the row of the FLADE fan
blades drivingly connected to the aft counter-rotatable fan having
radially inner airfoils.
BRIEF DESCRIPTION OF THE DRAWINGS
[0012] The foregoing aspects and other features of the invention
are explained in the following description, taken in connection
with the accompanying drawings where:
[0013] FIG. 1 is a schematical cross-sectional view illustration of
a first exemplary embodiment of a FLADE aircraft gas turbine engine
with chords of counter-rotatable fans having different stagger
angles at inner and outer airfoil cross-sections at a shroud
therebetween.
[0014] FIG. 2 is an enlarged cross-sectional view illustration of a
fan section of a more particular embodiment of the engine
illustrated in FIG. 1.
[0015] FIG. 3 is a perspective view illustration of the one of the
fans engine illustrated in FIG. 2.
[0016] FIG. 4 is a perspective view illustration taken in a
circumferential direction of a segment of the one of the fans
illustrated in FIG. 2.
[0017] FIG. 5 is a forward looking aft perspective view
illustration of the segment of the one of the fans illustrated in
FIG. 2.
[0018] FIG. 6 is a perspective view illustration of the segment of
the one of the fans illustrated in FIG. 2 normal to the inner
blade.
[0019] FIG. 7 is a schematical cross-sectional planform view
illustration of the inner and outer airfoil cross-sections and the
respective chords at the shroud for a ratio of outer airfoils to
inner airfoils of 1:1.
[0020] FIG. 8 is a schematical cross-sectional planform view
illustration of the inner and outer airfoil cross-sections and the
respective chords at the shroud for a ratio of outer airfoils to
inner airfoils of 2:1.
[0021] FIG. 9 is a schematical cross-sectional planform view
illustration of the inner and outer airfoil cross-sections and the
respective chords at the shroud for a ratio of outer airfoils to
inner airfoils of 3:1.
DETAILED DESCRIPTION OF THE INVENTION
[0022] Illustrated in FIGS. 1-3 is a FLADE counter-rotating fan
aircraft gas turbine engine 1 having a fan inlet 11 leading to
upstream and downstream or forward and aft counter-rotatable fans
130, 132. A circumferential row of fan inlet guide vanes 35 is
disposed between the fan inlet 11 and the forward counter-rotatable
fan 130. The FLADE fan assembly 60 includes a FLADE fan 2 having at
least one row of FLADE fan blades 5 disposed in a FLADE duct 3
through which FLADE airflow 80 is exhausted by the FLADE fan blades
5. The row of FLADE fan blades 5 is disposed radially outward of,
operably connected to, and driven by one of the forward or aft
counter-rotatable fans 130, 132. The FLADE fan blades 5 are
disposed axially aft and downstream of variable FLADE inlet guide
vanes 6. In FIG. 1, the aft fan 132 is illustrated as the FLADE fan
having the row of FLADE fan blades 5. The FLADE fan 2 is disposed
downstream of an annular FLADE inlet 8 to the FLADE duct 3. The
FLADE inlet 8 and the fan inlet 11 in combination generally form a
FLADE engine inlet 13 having a FLADE engine inlet area AI. The
FLADE airflow 80 may be used for cooling such as cooling of a
substantially hollow centerbody 72 or used for other purposes.
Optionally, a portion of the FLADE airflow 80 may be injected into
an exhaust flow 122 of the engine 1 through an aft FLADE variable
area bypass injector door 144.
[0023] Downstream and axially aft of the forward and aft
counter-rotatable fans 130, 132 is a core engine 18 having an
annular core engine inlet 17 and a generally axially extending axis
or centerline 12 generally extending forward 14 and aft 16. A fan
bypass duct 40 located downstream and axially aft of the forward
and aft counter-rotatable fans 130, 132 circumscribes the core
engine 18. The FLADE duct 3 circumscribes the forward and aft
counter-rotatable fans 130, 132 and the fan bypass duct 40.
[0024] One important criterion of inlet performance discussed is
the ram recovery factor. A good inlet must have air-handling
characteristics which are matched with the engine, as well as low
drag and good flow stability. During supersonic operation of the
engine, if AI is too small to handle the inlet airflow, the inlet
shock moves downstream of the inlet throat and pressure recovery
across the shock worsens and the exit corrected flow from the inlet
increases to satisfy the engine demand. If AI is too large, the
FLADE engine inlet 13 will supply more air than the engine can use
resulting in excess drag (spillage drag), because we must either
by-pass the excess air around the engine or "spill" it back out of
the inlet. Too much air or too little air is detrimental to
aircraft system performance. The FLADE fan 2 and the FLADE duct 3
are designed and operated to help manage the inlet airflow
delivered by the inlet to the fans.
[0025] The fan inlet 11 is sized to receive essentially full engine
airflow 15 of the engine at full power conditions with the FLADE
engine inlet 13 essentially closed off by closing the variable
FLADE inlet guide vanes 6. The engine is further designed and
operated to fully open the inlet of the flade duct at predetermined
part power flight conditions and essentially close it at full power
conditions such as take-off. The aft counter-rotatable fan 132 has
a single row of generally radially outwardly extending and
circumferentially spaced-apart second fan blades 32. The FLADE fan
blades 5 and the second fan blades 32 are separated by a rotatable
annular shroud 9 to which the FLADE fan blades 5 are mounted. The
forward counter-rotatable fan 130 has a single row of generally
radially outwardly extending and circumferentially spaced-apart
first fan blades 33. The FLADE fan blades 5 are primarily used to
flexibly match inlet airflow requirements.
[0026] The FLADE fan assemblies disclosed in the prior art have
radially inner fan blades and the radially outer blades or FLADE
fan blades nearly continuous in section properties at a transition
region of a flade fan shroud between the inner and outer blades.
The prior art also discloses an equal number of FLADE fan blades
and first or second fan blades, whichever they were attached to. In
some prior art FLADE fan assemblies, the radially outer blades or
FLADE fan blades are considered extensions of the radially inner
fan blades.
[0027] The second fan blades 32 and the FLADE fan blades 5 include
radially inner and outer airfoils 61, 62 circumferentially disposed
about the centerline 12 and extending radially outwardly from
radially inner and outer bases 111, 112 to radially inner and outer
tips 114, 116, respectively as illustrated in FIGS. 2-6. Note that
the FLADE fan assembly 60 illustrated herein is a single piece ring
and that a segment of the assembly or ring is illustrated in FIGS.
4-6 to further illustrate a contour of the shroud 9 and the inner
and outer airfoils 61, 62. Referring further to FIGS. 7, 8 and 9,
the inner and outer airfoils 61, 62 have inner and outer airfoil
cross-sections 118, 120 at the shroud 9. Inner and outer chords
124, 126 of the inner and outer airfoil cross-sections 118, 120
extend between inner and outer leading and trailing edges ILE, ITE
and OLE, OTE of the inner and outer airfoils 61, 62 respectively.
Inner and outer stagger angles 168, 170 are defined as angles
between or with respect to the inner and outer chords 124, 126 at
the shroud 9 and the centerline 12.
[0028] In order to optimize design and operation of at least one of
the counter-rotatable fans, illustrated herein as the downstream or
aft counter-rotatable fan 132 or the radially inner fan blade
illustrated herein as the second fan blade 32, and to take
advantage of the aerodynamic efficiency and lower rotational speed
offered by the counter-rotatable fans, the inner and outer stagger
angles 168, 170 are different. The FLADE fan blades 5 and the outer
airfoils 62 equal the second fan blades 32 and the inner airfoils
61 in the exemplary embodiment of the FLADE fan assembly 60
illustrated in FIG. 7. The FLADE fan blades 5 and the outer
airfoils 62 outnumber the second fan blades 32 and the inner
airfoils 61 in the exemplary embodiments of the FLADE fan assembly
60 illustrated in FIGS. 8 and 9. Different embodiments of the FLADE
fan assembly 60 may employ a ratio of the FLADE fan blades 5 and
the outer airfoils 62 to the second fan blades 32 and the inner
airfoils 61 in a range of 1.5:1 to 4:1. The ratio of FLADE fan
blades 5 and the outer airfoils 62 to the second fan blades 32 and
the inner airfoils 61 in the exemplary embodiment of the FLADE fan
assembly 60 as illustrated in FIGS. 3-6 and 8 is 2:1. The ratio of
FLADE fan blades 5 and the outer airfoils 62 to the second fan
blades 32 and the inner airfoils 61 in the exemplary embodiment of
the FLADE fan assembly 60 as illustrated in FIG. 7 is 1:1 and in
FIG. 9 is 3:1.
[0029] Linear radial load paths LP extend radially along radii R
from the centerline 12 and from a disk 25 of the aft
counter-rotatable fan 132 through the inner and outer airfoils 61,
62 and the rotating shroud 9 between them. The exemplary embodiment
of the FLADE fan assembly 60 illustrated in FIGS. 3-6 and 7 has two
outer airfoils 62 for each of the inner airfoils 61 while the
embodiment of the FLADE fan assembly 60 illustrated in FIG. 8 has
three outer airfoils 62 for each two of the inner airfoils 61. A
first portion 134 of the radially extending linear load paths LP
passing near or through the inner and outer leading edges ILE, OLE
and a second portion 136 of the radially extending linear load
paths LP passing near or through the inner and outer trailing edges
ITE, OTE.
[0030] The embodiment of the FLADE fan assembly 60 illustrated in
FIG. 8 has three outer airfoils 62 for each two of the inner
airfoils 61 and further includes a third portion 140 of the
radially extending linear load paths LP passing through inner and
outer points 148, 150 between the inner and outer trailing edges
ITE, OTE along the inner and outer chords 124, 126 respectively. In
general there may be multiple portions of the radially extending
linear load paths LP passing through inner and outer points 148,
150 between the inner and outer trailing edges ITE, OTE along the
inner and outer chords 124, 126 respectively. The inner and outer
points 148, 150 need not be near the inner and/or the outer
trailing edges ITE, OTE.
[0031] The exemplary FLADE fan assembly 60 may have one or more
circumferential rows of the radii R. The exemplary embodiments of
the FLADE fan assembly 60 illustrated in FIGS. 7, 8 and 9 include
at least a first circumferential row 152 of the radii R extending
radially outwardly from and normal to the centerline 12 and each
one of the radii R in the first circumferential row 152 intersects
one of the inner chords 124 and one of the outer chords 126. More
particularly, a second circumferential row 154 of the radii R
extends radially outwardly from and normal to the centerline 12,
the radii R in the first circumferential row 152 pass near or
through the inner and outer leading edges ILE, OLE, and the radii R
in the second circumferential row 154 pass near or through the
inner and outer trailing edges ITE, OTE. Illustrated in FIG. 8 is a
third circumferential row 156 of the radii R extending radially
outwardly from and normal to the centerline 12 wherein the radii R
in the third circumferential row 156 pass between the inner and
outer trailing edges ITE, OTE along the inner and outer chords 124,
126 respectively. In general, there may be multiple rows of the
radii R intersecting the inner and outer chords 124, 126
respectively. The rows of the radii R need not intersect the inner
and outer chords 124, 126 at or near the inner and/or the outer
trailing edges ITE, OTE.
[0032] The variable FLADE inlet guide vanes 6 which controls swirl
into the FLADE fan blades 5 and their outer airfoils 62, a contour
of the shroud 9, the inner chord 124 are designed to achieve a
desired angular and radial alignment between the inner and outer
chords 124, 126 of the inner and outer airfoil cross-sections 118,
120, the inner and outer stagger angles 168, 170, and the
intersection between the inner chords 124 and one of the outer
chords 126. This results in a unique mixed-flow flowpath
configuration at the radially inner tips 114 of the second fan
blades 32 which helps reduce shock losses of the second fan blades
32. The embodiment of the FLADE fan assembly 60 having different
inner and outer stagger angles 168, 170 as illustrated herein has a
particular application to the aft counter-rotatable fan 132.
[0033] Referring back to FIG. 1, the core engine 18 includes, in
downstream serial axial flow relationship, a core driven fan 37
having a row of core driven fan blades 36, a high pressure
compressor 20, a combustor 22, and a high pressure turbine 23
having a row of high pressure turbine blades 24. A high pressure
shaft 26 disposed coaxially about the centerline 12 of the engine 1
fixedly interconnects the high pressure compressor 20 and the high
pressure turbine blades 24. The core engine 18 is effective for
generating combustion gases. Pressurized air from the high pressure
compressor 20 is mixed with fuel in the combustor 22 and ignited,
thereby, generating combustion gases. Some work is extracted from
these gases by the high pressure turbine blades 24 which drives the
core driven fan 37 and the high pressure compressor 20. The high
pressure shaft 26 rotates the core driven fan 37 having a single
row of circumferentially spaced apart core driven fan blades 36
having generally radially outwardly located blade tip sections 38
separated from generally radially inwardly located blade hub
sections 39 by an annular fan shroud 108.
[0034] The combustion gases are discharged from the core engine 18
into counter-rotatable first and second low pressure turbines 19,
21 having first and second rows of low pressure turbine blades 28,
29, respectively. The second low pressure turbine 21 is drivingly
connected to the forward counter-rotatable fan 130 by a first low
pressure shaft 30, the combination or assembly being designated a
first low pressure spool 240. The first low pressure turbine 19 is
drivingly connected to the aft counter-rotatable fan 132 by a
second low pressure shaft 31, the combination or assembly being
designated a second low pressure spool 242. The high pressure
turbine 23 includes a row of high pressure turbine (HPT) nozzle
stator vanes 110 which directs flow from the combustor 22 to the
row of high pressure turbine blades 24.
[0035] Flow from the row of high pressure turbine blades 24 is then
directed into the counter-rotatable second and first low pressure
turbines 21 and 19 and the second and first rows of low pressure
turbine blades 29 and 28, respectively. The exemplary embodiment of
the engine 1 illustrated in FIGS. 1-2, includes a row of low
pressure stator vanes 66 between the second and first rows of low
pressure turbine blades 29 and 28. A variable throat area engine
nozzle 218 having a variable nozzle throat A8 is downstream and
axially aft of the counter-rotatable second low pressure turbine 21
and the fan bypass duct 40.
[0036] Referring to FIGS. 1 and 2, a first bypass inlet 42 to the
fan bypass duct 40 is disposed axially between the aft
counter-rotatable fan 132 and the annular core engine inlet 17 to
the core engine 18, thereby, providing two coaxial bypass flowpaths
into the fan bypass duct from the forward and aft counter-rotatable
fans 130, 132. The first fan blades 33 of the forward
counter-rotatable fan 130 and the second fan blades 32 of the aft
counter-rotatable fan 132 are radially disposed across a first fan
duct 138. The row of circumferentially spaced-apart fan inlet guide
vanes 35 is radially disposed across the first fan duct 138,
upstream and axially forward of the forward and aft
counter-rotatable fan 130, 132. The first fan duct 138 contains the
forward and aft counter-rotatable fans 130, 132 including the first
and second fan blades 33, 32 and the row of circumferentially
spaced-apart fan inlet guide vanes 35. The row of the core driven
fan blades 36 of the core driven fan 37 are radially disposed
across an annular second fan duct 142. The second fan duct 142
begins axially aft of the first bypass inlet 42 and is disposed
radially inwardly of the fan bypass duct 40. An annular first flow
splitter 45 is radially disposed between the first bypass inlet 42
and the second fan duct 142.
[0037] The full engine airflow 15 is split between the FLADE inlet
8 and the fan inlet 11. A fan airflow 50 passes through the fan
inlet 11 and then the forward and aft counter-rotatable fans 130,
132. A first bypass air portion 52 of the fan airflow 50 passes
through the first bypass inlet 42 of the fan bypass duct 40 when a
front variable area bypass injector (VABI) door 44 in the first
bypass inlet 42 is open and with the remaining air portion 54
passing through the core driven fan 37 and its row of core driven
fan blades 36. A row of circumferentially spaced-apart core driven
fan stator vanes 34 within the second fan duct 142 are disposed
axially between the row of second fan blades 32 and the core driven
fan blades 36 of the core driven fan 37. The row of the core driven
fan stator vanes 34 and the core driven fan blades 36 of the core
driven fan 37 are radially disposed across the second fan duct 142.
A vane shroud 106 divides the core driven fan stator vanes 34 into
radially vane hub sections 85 and vane tip sections 84,
respectively. The fan shroud 108 divides the core driven fan blades
36 into radially blade hub sections 39 and blade tip sections 38,
respectively.
[0038] A second bypass airflow portion 56 is directed through a fan
tip duct 146 across the vane tip sections 84 of the core driven fan
stator vanes 34 and across the blade tip sections 38 of the core
driven fan blades 36 into a second bypass inlet 46 of a second
bypass duct 58 to the fan bypass duct 40. An optional middle
variable area bypass injector (VABI) door 83 may be disposed at an
aft end of the second bypass duct 58 for modulating flow through
the second bypass inlet 46 to the fan bypass duct 40. An aft
variable area bypass injector (VABI) door 49 is disposed at an aft
end of the fan bypass duct 40 to mix bypass air 78 with core
discharge air 70.
[0039] The fan tip duct 146 includes the vane and fan shrouds 106,
108 and a second flow splitter 55 at a forward end of the vane
shroud 106. First and second varying means 91, 92 are provided for
independently varying flow areas of the vane hub and tip sections
85, 84, respectively. Exemplary first and second varying means 91,
92 include independently variable vane hub and tip sections 85, 84,
respectively (see U.S. Pat. No. 5,806,303). The independently
variable vane hub and tip sections 85, 84 designs may include the
entire vane hub and tip sections 85, 84 being independently
pivotable. Other possible designs are disclosed in U.S. Pat. Nos.
5,809,772 and 5,988,890.
[0040] Another embodiment of the independently variable vane hub
and tip sections 85, 84 includes pivotable trailing-edge hub and
tip flaps 86, 88 of the independently variable vane hub and tip
sections 85, 84 as illustrated in FIG. 1. The first and second
varying means 91, 92 can include independently pivoting flaps.
Alternative varying means for non-pivotable fan stator vane designs
include axially moving unison rings and those means known for
mechanical clearance control in jet engines (i.e., mechanically
moving circumferentially surrounding shroud segments radially
towards and away from a row of rotor blade tips to maintain a
constant clearance despite different rates of thermal expansion and
contraction). Additional such varying means for non-pivotable, fan
stator vane designs include those known for extending and
retracting wing flaps on airplanes and the like.
[0041] Exemplary first and second varying means 91, 92, illustrated
in FIG. 1 include an inner shaft 94 coaxially disposed within an
outer shaft 96. The inner shaft 94 is rotated by a first lever arm
98 actuated by a first unison ring 100. The outer shaft 96 is
rotated by a second lever arm 102 actuated by a second unison ring
104. The inner shaft 94 is attached to the pivotable trailing edge
hub flap 86 of the vane hub section 85 of the fan stator vane 34.
The outer shaft 96 is attached to the pivotable trailing edge tip
flap 88 of the vane tip section 84 of the fan stator vane 34. Note
that the lever arms 98, 102 and the unison rings 100, 104 are all
disposed radially outward of the fan stator vanes 34.
[0042] The forward and aft counter-rotatable fans 130, 132 in
counter-rotating fan engines allows the elimination of a row of
stator vanes between the counter-rotatable fans in the fan section
of the engine and also help to minimize the number of nozzles or
vanes in the turbine. The savings in weight and cost due to the
removal of the fan stator vanes is traded against the complexity of
adding a third spool, namely one of the forward and aft
counter-rotating low pressure spools. Counter-rotating fan engines
typically have a wheel speed of the aft counter-rotatable fan 132
that is somewhat lower than that of the forward counter-rotatable
fan 130. This is one reason for selecting the aft counter-rotatable
fan 132 upon which to mount the row of FLADE fan blades 5. An
elevated relative Mach number into the aft counter-rotatable fan
132 is the reason for its lower wheel speed and it is a result of
the counter-swirl imparted by the forward counter-rotatable fan
130. The lower wheel speed of the aft counter-rotatable fan 132
suggests a reduced work fraction on it to equalize the net fan
rotor torque. In this manner, the exit swirl from the aft
counter-rotatable fan 132 is sufficiently small so no downstream
straightening vanes are required. One exemplary speed ratio of the
aft counter-rotatable fan 132 to the forward counter-rotatable fan
130, (speed of rotor 2/speed of rotor 1), is 0.75 which also the
work ratio of the two fans. The resulting work split is 57.5% for
the forward counter-rotatable fan 130 and the remaining 42.5% for
the aft counter-rotatable fan 132. Current studies suggest that
energy requirements of the row of FLADE fan blades 5 is in a range
of 15 to 30 percent of the overall fan energy.
[0043] One problem with counter-rotatable fans is an area ratio
requirement across the first low pressure turbine 19. Prudent
design practice suggests little or no outward slope over the
turbine rotor to lessen turbine blade tip clearance migrations with
the axial migration of the turbine rotor. Design practice also
constrains turbine blade hub slope to less that about 30 degrees to
avoid excessive aerodynamic loss in this region. It is desirable to
avoid first low pressure turbines having rotor pressure ratios in
excess of about 1.45. Turbines rotor pressure ratio is defined
turbine blade inlet pressure divided by turbine blade exit
pressure. Prior counter-rotatable fan engine designs indicate that
first low pressure turbines have pressure ratios of about 1.9. This
is far more than what is desirable.
[0044] The total work on the second low pressure spool 242 is the
sum of the work performed by the aft counter-rotatable fan 132 plus
the work performed by have the FLADE fan blades 5. The total work
extracted by the first low pressure turbine 19, which is drivingly
connected to the aft counter-rotatable fan 132, requires a first
low pressure turbine 19 pressure ratio well in excess of the above
noted limit for a no turbine nozzle configuration. A solution to
this problem is to reduce the work requirement of the aft
counter-rotatable fan 132 to a point consistent with a first low
pressure turbine 19 pressure ratio of about 1.45. The reduced work
of the aft counter-rotatable fan 132 is then added to the work
required by forward counter-rotatable fan 130, thereby restoring
the total fan work.
[0045] Adequate fan stall margin must be retained with the revised
stage pressure ratio requirements. Rotor speeds of the forward and
aft counter-rotatable fans 130, 132 are determined by their
respective pressure ratio requirements. The rotor speed of the aft
counter-rotatable fan 132 is determined by its pressure ratio
requirement or alternatively by the pressure ratio requirement of
the FLADE fan blades 5. The resulting work ratio for the aft
counter-rotatable fan 132 in the engine illustrated in FIGS. 1-3 is
about 0.43 and its speed ratio is about 0.73.
[0046] The flade airflow 80 may be modulated using the variable
FLADE inlet guide vanes 6 to provide maximum engine airflow
capability at take-off operating conditions for noise abatement or
for engine-inlet airflow matching during flight. At supersonic
cruise conditions the flade airflow may be reduced to its least
energy absorbing airflow to permit the highest attainable specific
thrust. The flade airflow modulation may alter the work requirement
of the first low pressure turbine 19 of the second low pressure
spool 242. However, the first low pressure turbine 19 and its first
row of low pressure turbine blades 28 is nested between the row of
high pressure turbine blades 24 of the high pressure turbine 23 and
the second low pressure turbine 21 and its second row of low
pressure turbine blades 29.
[0047] The first low pressure turbine 19 inlet flow function is
expected to remain relatively constant over its steady state
operating space. The second low pressure turbine 21 inlet flow
function is also expected to remain relatively constant over its
steady state operating regime. Accordingly, the pressure ratio of
the first low pressure turbine 19 is expected to remain relatively
constant. At constant pressure ratio the work output of the first
low pressure turbine 19 will remain relatively constant. This
constant work output of the first low pressure turbine 19 coupled
with the reduced work input requirement of the first low pressure
spool 240, due to closure of the variable FLADE inlet guide vanes 6
and the row of FLADE fan blades 5 would create a torque imbalance
and cause an acceleration of the low pressure spool 240. The
pressure ratio across the first low pressure turbine 19 must be
modulated to prevent this excess torque. The modulation is
accomplished by varying the row of variable low pressure stator
vanes 66 between the first and second rows of low pressure turbine
blades 28, 29 to adjust inlet flow to the second row of low
pressure turbine blades 29. A variable throat area A8 helps to
avoid over extraction by the first low pressure turbine 19.
[0048] While there have been described herein what are considered
to be preferred and exemplary embodiments of the present invention,
other modifications of the invention shall be apparent to those
skilled in the art from the teachings herein and, it is therefore,
desired to be secured in the appended claims all such modifications
as fall within the true spirit and scope of the invention.
Accordingly, what is desired to be secured by Letters Patent of the
United States is the invention as defined and differentiated in the
following claims.
* * * * *