U.S. patent application number 12/886807 was filed with the patent office on 2012-03-22 for gas turbine engine with bleed duct for minimum reduction of bleed flow and minimum rejection of hail during hail ingestion events.
Invention is credited to Anthony R. Bifulco, Justin R. Urban.
Application Number | 20120070271 12/886807 |
Document ID | / |
Family ID | 44785427 |
Filed Date | 2012-03-22 |
United States Patent
Application |
20120070271 |
Kind Code |
A1 |
Urban; Justin R. ; et
al. |
March 22, 2012 |
GAS TURBINE ENGINE WITH BLEED DUCT FOR MINIMUM REDUCTION OF BLEED
FLOW AND MINIMUM REJECTION OF HAIL DURING HAIL INGESTION EVENTS
Abstract
A gas turbine engine includes a bleed structure which includes a
forward wall and a rear structural wall to define a deposit space
downstream of the bleed structure for a hail event of a
predetermined duration.
Inventors: |
Urban; Justin R.; (Tolland,
CT) ; Bifulco; Anthony R.; (Ellington, CT) |
Family ID: |
44785427 |
Appl. No.: |
12/886807 |
Filed: |
September 21, 2010 |
Current U.S.
Class: |
415/145 |
Current CPC
Class: |
F02K 3/06 20130101; Y02T
50/60 20130101; F01D 17/105 20130101; F02C 7/052 20130101; F02C
6/08 20130101; F02C 6/18 20130101; F05D 2260/607 20130101; F04D
29/545 20130101; Y02T 50/675 20130101; F05D 2260/606 20130101; Y10T
29/49231 20150115; F02C 7/05 20130101; F04D 27/023 20130101; F02C
7/047 20130101 |
Class at
Publication: |
415/145 |
International
Class: |
F01D 17/00 20060101
F01D017/00 |
Claims
1. A gas turbine engine comprising: a bleed structure which
includes a forward wall; and a rear structural wall to define a
deposit space downstream of said bleed structure.
2. The gas turbine engine as recited in claim 1, further comprising
a bleed duct aft lip located aft of said forward wall, said bleed
duct aft lip extends in a radial direction less than said forward
wall.
3. The gas turbine engine as recited in claim 1, wherein said rear
structural wall forms a portion of a core case section.
4. The gas turbine engine as recited in claim 1, further comprising
a circumferentially intermittent aft wall aft of said forward wall
to define said bleed duct.
5. The gas turbine engine as recited in claim 1, wherein said rear
structural wall backstops the bleed structure.
6. The gas turbine engine as recited in claim 1, further comprising
a split duct particle separation structure aft of said forward wall
which defines a bleed airflow path and a particle path aft of said
bleed airflow path.
7. The gas turbine engine as recited in claim 6, wherein said split
duct particle separation structure includes a split wall and a
particle separation wall.
8. The gas turbine engine as recited in claim 7, wherein said split
wall is radially outboard of said particle separation wall.
9. The gas turbine engine as recited in claim 1, wherein said rear
structural wall at least partially forms the bleed duct aft
wall.
10. The gas turbine engine as recited in claim 9, wherein said rear
structural wall separates a first engine section from a second
engine section.
11. The gas turbine engine as recited in claim 9, wherein said rear
structural wall is directly adjacent to a high pressure compressor
case.
12. The gas turbine engine as recited in claim 1, wherein said
bleed structure defines a station 2.5 bleed duct.
13. The gas turbine engine as recited in claim 1, wherein said
deposit space is sized for a hail event of a predetermined
duration.
14. A gas turbine engine comprising: a bleed structure which
includes a forward wall and an aft wall; and a fluid plenum at
least partially formed by said aft wall to receive a heated
fluid.
15. The gas turbine engine as recited in claim 14, wherein said
bleed structure defines a station 2.5 bleed duct.
16. The gas turbine engine as recited in claim 14, wherein said
heated fluid is oil.
17. The gas turbine engine as recited in claim 14, wherein said
heated fluid is air.
18. A method to minimize the formation of hail in a bleed passage
of a gas turbine engine comprising: defining a deposit space
downstream of a bleed structure for a hail event of a predetermined
duration.
19. A method as recited in claim 18, defining an intermittent aft
wall upstream of the deposit space.
20. A method as recited in claim 18, locating a rear structural
wall to at least partially form a bleed duct aft wall.
Description
BACKGROUND
[0001] The present disclosure relates to gas turbine engines;
particularly bleed flow handling for gas turbine engines.
[0002] In aircraft gas turbine engines, air is directed through
multiple stage compressors. As the air passes through each
successive compressor stage, the pressure of the air is increased.
Under certain conditions, such as when the engine is operating at
off design conditions, interstage bleed through various bleed ducts
is utilized to rematch the compressor stages. Typically, a station
2.5 bleed duct is also utilized to remove hail ice, ice crystals,
and accreted ice in flight.
SUMMARY
[0003] A gas turbine engine according to an exemplary aspect of the
present disclosure includes a bleed structure with a forward wall
and a rear structural wall to define a deposit space downstream of
the bleed structure for a hail event of a predetermined
duration.
[0004] A gas turbine engine according to an exemplary aspect of the
present disclosure includes a bleed structure with a forward wall
and an aft wall and a fluid plenum at least partially formed by the
aft wall to receive a heated fluid.
[0005] A method to minimize the formation of hail in a bleed
passage of a gas turbine engine according to an exemplary aspect of
the present disclosure includes defining a deposit space downstream
of a bleed structure for a hail event of a predetermined
duration.
BRIEF DESCRIPTION OF THE DRAWINGS
[0006] Various features will become apparent to those skilled in
the art from the following detailed description of the disclosed
non-limiting embodiment. The drawings that accompany the detailed
description can be briefly described as follows:
[0007] FIG. 1 is a general sectional view through a gas turbine
engine along the engine longitudinal axis;
[0008] FIG. 2 is an expanded side sectional view through a gas
turbine engine illustrating a 2.5 bleed structure;
[0009] FIG. 3 is an expanded sectional view of a RELATED ART 2.5
bleed structure;
[0010] FIG. 4 is an expanded sectional view of a 2.5 bleed
structure in accords with one embodiment of the present
disclosure;
[0011] FIG. 5 is an expanded sectional view of a 2.5 bleed
structure in accords with another embodiment of the present
disclosure;
[0012] FIG. 6 is a perspective partial sectional view of a
circumferentially intermittent aft wall of the 2.5 bleed structure
of FIG. 5;
[0013] FIG. 7 is an expanded sectional view of a 2.5 bleed
structure in accords with another embodiment of the present
disclosure;
[0014] FIG. 8 is an expanded sectional view of a 2.5 bleed
structure in accords with another embodiment of the present
disclosure;
[0015] FIG. 9 is a partial sectional view of a plenum of the 2.5
bleed structure of FIG. 8; and
[0016] FIG. 10 is an expanded sectional view of a 2.5 bleed
structure in accords with another embodiment of the present
disclosure.
DETAILED DESCRIPTION
[0017] FIG. 1 illustrates a general partial fragmentary schematic
view of a gas turbofan engine 10 suspended from an engine pylon 12
within an engine nacelle assembly N as is typical of an aircraft
designed for subsonic operation. While a two spool high bypass
turbofan engine with a geared architecture is schematically
illustrated in the disclosed non-limiting embodiment, it should be
understood that the disclosure is applicable to other gas turbine
engine configurations.
[0018] The turbofan engine 10 includes a core engine within a core
nacelle C that houses a low spool 14 and high spool 24. The low
spool 14 includes a low pressure compressor 16 and low pressure
turbine 18. The low spool 14 drives a fan section 20 connected to
the low spool 14 either directly or through a geared architecture
25. The high spool 24 includes a high pressure compressor 26 and
high pressure turbine 28. A combustor 30 is arranged between the
high pressure compressor 26 and high pressure turbine 28. The low
and high spools 14, 24 rotate about an engine axis of rotation
A.
[0019] Airflow enters the fan nacelle F which at least partially
surrounds the core nacelle C. The fan section 20 communicates
airflow into the core nacelle C to the low pressure compressor 16
and the high pressure compressor 26. Core airflow compressed by the
low pressure compressor 16 and the high pressure compressor 26 is
mixed with the fuel in the combustor 30, is ignited, and burned.
The resultant high pressure combustor products are expanded through
the high pressure turbine 28 and low pressure turbine 18. The
turbines 28, 18 are rotationally coupled to the compressors 26, 16
respectively to drive the compressors 26, 16 in response to the
expansion of the combustor product. The low pressure turbine 18
also drives the fan section 20 to communicate a bypass flow. A core
engine exhaust exits the core nacelle C through a core nozzle 43
defined between the core nacelle C and a tail cone 33.
[0020] With reference to FIG. 2, engine static structure 44
generally has sub-structures which may include a case structure
often referred to as the engine backbone. The fan section 20
includes a fan rotor 32 with a plurality of circumferentially
spaced radially outwardly extending fan blades 34. The fan blades
34 are surrounded by a fan case structure 44F. The core case
structure 44C is secured to the fan case structure 44F through a
multiple of circumferentially spaced radially extending fan exit
guide vanes (FEGVs) 40 which radially span a core case structure
44C and the fan case structure 44F defined about the engine axis
A.
[0021] A bleed structure 62 such as a 2.5 bleed duct structure is
typically located just forward of a rear structural wall 64 of the
core case structure 44C to direct core airflow compressed by the
low pressure compressor 16 selectively out into the bypass flow
stream through a bleed valve (not shown). It should be understood
that the bleed structure 62 may be of various bleed duct and bleed
door configurations as generally understood. The rear structural
wall 64 is located inboard of the FEGVs 40 to at least partially
provide support therefore. It should be understood that the shape
and configuration of the engine static structure 44 and rear
structural wall 64 may be of various forms.
[0022] Applicant has determined that hail event blockage within the
Station 2.5 bleed system usually takes place upon a bleed duct aft
wall W of a bleed duct structure D (RELATED ART; FIG. 3). Thus,
modification of the bleed duct aft wall W such that the flow area
will not be reduced when hail ice enters the station 2.5 bleed
system will solve the issue of bleed flow reduction and hail
extraction efficiency reduction during hail events.
[0023] With Reference to FIG. 4, a bleed structure 62A according to
one disclosed non-limiting embodiment essentially deletes the bleed
duct aft wall W but maintains a forward wall 66 and a bleed duct
aft lip 68. The forward wall 66 and the bleed duct aft lip 68
thereby defines the flow path for the 2.5 bleed structure 62A. The
forward wall 66 maintains the aerodynamic properties of the bleed
duct structure via a forward wall coanda effect. The hail ice
buildup base may also be minimized through optimization of the
length of the bleed duct aft lip 68 so as to increases shedding
frequency therefrom to direct the ice toward the rear structural
wall 64 which backstops the bleed structure 62A and provides the
necessary deposit space for a hail event of a predetermined
duration such as 30 seconds. That is, the bleed structure 62 is
defined by the forward wall 66, the lip 68 and the rear structural
wall 64 of the core case structure 44C rather than a specifically
defined duct (RELATED ART; FIG. 3). Ribs 70 or other static
structure may alternatively or additionally be utilized to provide
the desired structural support within the core case structure 44C
for the low pressure compressor vanes 16V which may have been
hereto for provided by the rear wall.
[0024] With reference to FIG. 5, a bleed structure 62B according to
another disclosed non-limiting embodiment, provides a
circumferentially intermittent bleed duct aft wall 72. That is, the
bleed duct aft wall 72 includes circumferential openings to provide
a tooth like structure which is also backstopped by the rear
structural wall 64 (FIG. 6).
[0025] With reference to FIG. 7, a bleed structure 62C according to
another disclosed non-limiting embodiment, includes a split duct
particle separation structure 74 which defines a bleed airflow path
76 and a particle path 78 aft of the bleed airflow path 76. A split
wall 80 separates the bleed airflow path 76 and the particle path
78. The split wall 80 and a particle separation wall 82 are
radially arranged relative the core flow path to separate the hail
from the airflow in a manner such as an inlet particle separator.
That is, the split wall 80 may be radially inboard of the particle
separation wall 82 such that hail is directed through the particle
path 78.
[0026] With reference to FIG. 8, a bleed structure 62D according to
another disclosed non-limiting embodiment, provides a plenum 84 at
least partially formed by an bleed duct aft wall 86. The plenum 84
circulates a fluid such as warmer bleed air or oil as a heating
fluid to inhibit the accumulation or formation of hail/ice in the
2.5 bleed passage. As the geared architecture 25 (FIG. 1) receives
warm air and oil which is communicated forward of the LPC, either
or both of these fluids is diverted into the plenum 84. It should
be understood that other non-geared architecture designs may obtain
the fluid from other sources. Passage may be provided through
structural walls via local pass through conduits 86. In the
disclosed non-limiting embodiment, the fluid enters at one or more
locations and is pumped into and around the plenum 84 (FIG. 9) then
communicated to a destination such as the geared architecture
25.
[0027] With reference to FIG. 10, a bleed structure 62E according
to another disclosed non-limiting embodiment locates the rear
structural wall 64' of the core case structure 44C which at least
partially forms the bleed duct bleed duct aft wall 88. The rear
structural wall 64' is directly adjacent to, for example, high
pressure compressor case module 90 which generally operates at a
higher temperature and/or has greater mass and thereby heats the
bleed duct bleed duct aft wall 88 to inhibit the accumulation or
formation of hail/ice in the 2.5 bleed passage.
[0028] If no significant degree of ice clogging were to take place
due to specific design features, then a rig test is not required to
determine the level of clogging. The bleed duct structures
disclosed herein may enable elimination of hail ingestion
certification rigs and allow turbofan engines to digest more hail
than current designs. The bleed duct structures disclosed herein
will also provide an increased margin during hail ingestion events
and may decrease overall weight. Furthermore, regulations may
change over time to require increased hail ingestion and the bleed
duct structures disclosed herein will readily accommodate such an
increase.
[0029] The foregoing description is exemplary rather than defined
by the limitations within. Many modifications and variations of the
present invention are possible in light of the above teachings. The
disclosed embodiments of this invention have been disclosed,
however, one of ordinary skill in the art would recognize that
certain modifications would come within the scope of this
invention. It is, therefore, to be understood that within the scope
of the appended claims, the invention may be practiced otherwise
than as specifically described. For that reason the following
claims should be studied to determine the true scope and content of
this invention.
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