U.S. patent application number 11/232508 was filed with the patent office on 2007-03-22 for turbine engine nozzle.
This patent application is currently assigned to United Technologies Corporation. Invention is credited to Curtis C. Cowan, Debora F. Kehret.
Application Number | 20070062199 11/232508 |
Document ID | / |
Family ID | 37669660 |
Filed Date | 2007-03-22 |
United States Patent
Application |
20070062199 |
Kind Code |
A1 |
Cowan; Curtis C. ; et
al. |
March 22, 2007 |
Turbine engine nozzle
Abstract
A turbine engine nozzle assembly has an upstream flap assembly
having a main flap and a liner, a cooling passageway formed between
the main flap and liner. A downstream flap is pivotally coupled to
the upstream flap assembly for relative rotation about a hinge
axis. The liner has a trailing end spaced upstream from a trailing
end of the main flap by at least 40% of a length of the main
flap.
Inventors: |
Cowan; Curtis C.; (East
Hampton, CT) ; Kehret; Debora F.; (Manchester,
CT) |
Correspondence
Address: |
BACHMAN & LAPOINTE, P.C. (P&W)
900 CHAPEL STREET
SUITE 1201
NEW HAVEN
CT
06510-2802
US
|
Assignee: |
United Technologies
Corporation
|
Family ID: |
37669660 |
Appl. No.: |
11/232508 |
Filed: |
September 22, 2005 |
Current U.S.
Class: |
60/770 ;
239/265.19 |
Current CPC
Class: |
F02K 1/38 20130101; Y02T
50/60 20130101; Y02T 50/671 20130101; F02K 1/1292 20130101; F02K
1/1223 20130101 |
Class at
Publication: |
060/770 ;
239/265.19 |
International
Class: |
B63H 11/10 20060101
B63H011/10 |
Goverment Interests
U.S. GOVERNMENT RIGHTS
[0001] The invention was made with U.S. Government support under
contract no. N00019-02-C-3003 awarded by the U.S. Navy. The U.S.
Government has certain rights in the invention.
Claims
1. A turbine engine nozzle subassembly comprising: an upstream flap
assembly having a main flap and a liner, a cooling passageway
formed between the main flap and liner; a downstream flap pivotally
coupled to the upstream flap for relative rotation about a hinge
axis; and an actuator linkage coupled to at least one of the
upstream flap and the downstream flap for actuating the upstream
and downstream flaps between a plurality of throat area conditions,
wherein: the liner has a trailing end spaced upstream from a
trailing end of the main flap by at least 40% of a length of the
main flap.
2. The subassembly of claim 1 further comprising: an external flap
pivotally coupled to the downstream flap and to an environmental
structure so that a span between respective coupling locations with
said downstream flap and environmental structure is extensible and
contractable responsive to aerodynamic forces.
3. The subassembly of claim 1 wherein the liner comprises: a liner
body; and a liner mounting bracket secured to the liner body and to
the main flap.
4. The subassembly of claim 3 wherein: the liner body comprises an
Nb-based sheet and a Ni-based superalloy backing element.
5. The subassembly of claim 1 wherein: the liner trailing end is
spaced upstream from the main flap trailing end of the main flap by
70-80% of the length of the main flap.
6. The subassembly of claim 1 wherein: the liner has a length of
15-50% of the length of the main flap.
7. The subassembly of claim 1 wherein: the liner has a length of
20-30% of the length of the main flap.
8. A turbine engine nozzle comprising: a static structure; a
plurality of flap subassemblies comprising: an upstream main flap
pivotally coupled to the static structure for relative rotation
about an axis essentially fixed relative to the static structure;
and a downstream flap pivotally coupled to the upstream flap for
relative rotation about a hinge axis; and a liner along the
upstream main flaps and forming a generally annular cooling air
passageway, the cooling passageway having an outlet spaced upstream
of a downstream end of the main flaps by a longitudinal distance of
at least 40% of a longitudinal length of the upstream main
flaps.
9. The nozzle of claim 8 wherein: the plurality of flap
subassemblies are axisymmetrically arranged about an engine
centerline; said articulation is simultaneous for each of the flap
subassemblies; and each of the plurality of flap subassemblies
further comprises an external flap pivotally coupled to the
downstream flap.
10. The nozzle of claim 8 wherein the liner comprises a
circumferential array of: a plurality of first members, each
mounted to an associated one of the main flaps; and a plurality of
second members, each between an associated pair of the first
members and mounted to an associated convergent seal.
11. A turbine engine nozzle comprising: a static structure; a
convergent section comprising: a circumferential array of first
flaps, each pivotally coupled to the static structure; a
circumferential array of first seals, alternatingly interspersed
with the first flaps; and a liner assembly; a divergent section
comprising: a circumferential array of second flaps, each pivotally
coupled to an associated one of the first flaps; and a
circumferential array of second seals, alternatingly interspersed
with the second flaps, wherein the liner has an outlet spaced
upstream of a downstream end of the main flaps by a longitudinal
distance of essentially at least 40% of a longitudinal length of
the convergent section.
12. A gas turbine engine nozzle convergent section liner member
comprising: a panel having: an inboard surface; an outboard
surface; a leading end; a trailing end first and second lateral
ends; a length between the leading end and the trailing end; and a
lateral span between the first and second lateral ends, wherein:
the lateral span is greater than the length.
13. The liner member of claim 12 wherein: the length is 40-60% of
the lateral span.
14. The liner member of claim 12 wherein: the panel has: a
generally planar central portion; and means along the first and
second lateral edges for interfitting with complementary features
of a complementary panel.
15. The liner member of claim 12 further comprising: a mounting
bracket secured to the panel and extending from the outboard
surface and having: a central web essentially parallel and spaced
apart from a central portion of the panel and having a bolting
aperture; and first and second lateral webs extending toward the
panel from first and second edges of the central web.
16. The liner member of claim 12 wherein: the panel comprises a
liner sheet, a backing sheet along only an upstream portion of the
liner sheet, and a deflector; a plurality of rivets securing the
liner sheet, backing sheet, and deflector; and a pair of welds
secure the mounting bracket to the liner sheet.
17. A method for retrofitting a turbine engine or reengineering a
turbine engine configuration which engine or configuration has or
has previously had a first nozzle subassembly having a convergent
flap, a divergent flap, an external flap, and an actuation linkage
coupled to the convergent flap, the method comprising: replacing a
first liner member of the convergent flap with a second liner
member, the second liner member having a downstream end positioned
upstream from a former position of a downstream end of the first
liner member by at least 10% of a length of the convergent
flap.
18. The method of claim 17 wherein: said second liner member
provides a higher coolant-to-gas .rho.v ratio than was provided by
the first liner member.
19. The method of claim 17 wherein: said second liner member
comprises a liner sheet and a mounting bracket welded to the liner
sheet.
20. The method of claim 17 wherein: a plurality of such first liner
members of a circumferential array of such first nozzle
subassemblies are replaced with a plurality of such second liner
members.
Description
BACKGROUND OF THE INVENTION
[0002] The invention relates to turbine engines. More particularly,
the invention relates to variable throat turbine engine exhaust
nozzles.
[0003] There is well developed field in turbine engine exhaust
nozzles. A number of nozzle configurations involve pairs of
relatively hinged flaps: a convergent flap upstream; and a
divergent flap downstream. Axisymmetric nozzles may feature a
circular array of such flap pairs. Exemplary nozzles are shown in
U.S. Pat. Nos. 3,730,436, 5,797,544, and 6,398,129 and United
Kingdom patent application GB2404222 A.
SUMMARY OF THE INVENTION
[0004] Accordingly, one aspect of the invention involves a turbine
engine nozzle subassembly. An upstream flap assembly has a main
flap and a liner. A cooling passageway formed between the main flap
and liner. A downstream flap is pivotally coupled to the upstream
flap assembly for relative rotation about a hinge axis. The liner
has a trailing end spaced upstream from a trailing end of the main
flap by at least 40% of a length of the main flap.
[0005] The details of one or more embodiments of the invention are
set forth in the accompanying drawings and the description below.
Other features, objects, and advantages of the invention will be
apparent from the description and drawings, and from the
claims.
BRIEF DESCRIPTION OF THE DRAWINGS
[0006] FIG. 1 is a cutaway longitudinal view of a prior art turbine
engine nozzle.
[0007] FIG. 2 is a partial view of assembled flaps of the nozzle of
FIG. 1
[0008] FIG. 3 is cutaway longitudinal view of a modified turbine
engine nozzle.
[0009] FIG. 4 is an enlarged view of a convergent flap of the
nozzle of FIG. 3.
[0010] FIG. 5 is a partial view of assembled flaps of the nozzle of
FIG. 3.
[0011] FIG. 6 is an exploded view of a convergent flap liner member
of the nozzle of FIG. 3.
[0012] FIG. 7 is an oblique outboard view of the liner member of
FIG. 6.
[0013] FIG. 8 is an oblique inboard view of the liner member of
FIG. 6.
[0014] FIG. 9 is an oblique inboard view of a convergent seal liner
member of the nozzle of FIG. 3
[0015] FIG. 10 is an oblique outboard view of the convergent seal
liner number of FIG. 9.
[0016] FIG. 11 is a transverse sectional view of the nozzle of FIG.
3.
[0017] FIG. 12 is a schematic longitudinal sectional view of the
nozzle of FIG. 3.
[0018] FIG. 13 is a graph of gas mass flux for cooling and main
flows against liner length.
[0019] FIG. 14 is a graph relating blowing ratios of baseline and
modified nozzles.
[0020] Like reference numbers and designations in the various
drawings indicate like elements.
DETAILED DESCRIPTION
[0021] FIG. 1 shows a prior art turbine engine nozzle 20. The
exemplary nozzle is drawn from United Kingdom patent application
GB2404222 A (the disclosure of which is incorporated by reference
herein as if set forth at length) and may serve as a baseline for
modifications discussed below. The exemplary nozzle 20 comprises an
axisymmetric circular array of convergent/divergent flap pairs
about a nozzle axis or centerline 500. A given flap pair has a
convergent flap 22 upstream/forward extending from an upstream end
23 to a downstream end 24 and a divergent flap 26 downstream/aft
extending from an upstream end 27 to a downstream end 28. The flaps
are hinged relative to each other by a hinge mechanism 30 for
relative movement about a hinge axis 502 proximate the convergent
flap downstream end and divergent flap upstream end. A concentric
circular array of seals may be interspersed with the flap pair
array. Exemplary seals (FIG. 2) have respective convergent members
32 and divergent members 33 between adjacent convergent flaps and
divergent flaps, respectively. The inboard surface of the divergent
flap 26 has a longitudinally convex surface portion 40 (FIG. 1)
near its upstream end for forming an aerodynamic throat (i.e., the
location of smallest passageway cross-section) of the nozzle of
instantaneous throat radius R.sub.T and an essentially
longitudinally straight portion 42 extending aft therefrom toward
the downstream end for forming an exhaust outlet of instantaneous
outlet radius R.sub.O. For each convergent/divergent flap pair, the
nozzle further includes an external flap 50, the outboard surface
52 of which forms an exterior contour of the nozzle exposed to
external airflow passing around the aircraft fuselage.
[0022] FIG. 1 further shows a nozzle static ring structure 60 for
mounting the nozzle to the engine, aircraft fuselage, or other
environmental structure. Proximate the upstream end 23 of the
convergent flap 22, a hinge structure pivotally couples the
convergent flap to the static ring structure 60 for relative
rotation about a fixed transverse axis 503. A synchronization ring
62 is mounted between inboard and outboard aft portions 64 and 66
of the static ring structure and may be longitudinally reciprocated
by actuators (e.g., pneumatic or hydraulic actuators-not shown). In
the condition of FIG. 1, the synchronization ring is at a
forwardmost/upstreammost position. The synchronization ring is
coupled to each flap pair by an associated linkage 70. Each linkage
70 includes a central bell crank 72 pivotally coupled by a hinge
mechanism to a bell crank ground point 74 at the trailing edge of
the static ring structure inboard portion 64 for relative rotation
about a fixed transverse axis 504. To drive rotation of the bell
crank through its range of rotation about the axis 504, the bell
crank is coupled to the synchronization ring by an associated
H-link 76. A forward end of the H-link is pivotally coupled to the
synchronization ring by a hinge mechanism for relative rotation
about a transverse axis 506 which shifts longitudinally with the
synchronization ring. An aft end of the H-link is pivotally coupled
to the bell crank by a hinge mechanism for relative rotation about
a transverse axis 508 which moves along a circular path segment
centered about the axis 504 in response to linear translation of
the axis 506. Thus, as viewed in FIG. 1, a rearward shift of the
synchronization ring produces a clockwise rotation of the bell
crank about the axis 504. Rotation of the bell crank is transferred
to articulation of the associated flap pair by an associated pair
of transfer links 78. Forward/upstream ends of each pair of
transfer links are pivotally coupled to the bell crank for relative
rotation about a transverse axis 510 which also moves along a
circular path segment centered about the axis 504 in response to
linear translation of the axis 506. Aft/downstream ends of the
transfer links are pivotally coupled to the divergent flap 26 for
relative rotation about a transverse axis 512. As discussed below,
in the exemplary embodiment movement of the axis 512 is not
entirely dictated by the rotation of the bell crank and associated
static ring translation. Rather, it may be influenced by other
forces, namely aerodynamic forces arising from relative pressures
internal and external to the nozzle. In exemplary embodiments, the
axis 512 falls aft of the axis 502 and along a forward half of the
span between upstream and downstream ends of the divergent flap.
More narrowly, it falls along a forward third, and, in the
illustrated embodiment, approximately in between about the first 5%
and 15% of such span.
[0023] In the exemplary embodiment, the external flap 50 has a
forward end 90 pivotally coupled by a hinge mechanism to the static
structure outboard portion 66 for relative rotation about a fixed
transverse axis 520. Proximate its downstream end 92, the external
flap is pivotally coupled by a hinge mechanism to the divergent
flap 26 (slightly more forward of its downstream end 28) for
relative rotation about a transverse axis 522. The external flap is
configured so that the span between the axes 520 and 522 is
extensible and contractible such as by having an upstream link 94
telescopically mounted relative to a main body portion 96 of the
external flap and coupling the external flap to the static ring
structure. The extensibility/contractability may have a limited
range. For a further limitation on that range, a secondary link or
mode strut 100 is provided having a forward end portion 102
pivotally coupled to the static ring structure for relative
rotation about a fixed transverse axis 524 which may be close to
the axis 520. If the axes 520 and 524 are coincident, it may be
advantageous to drill one hole through all pivot points for low
cost. However, if the width of the external flap 50 is such that
the main body portion 96 on either circumferential side of the flap
are substantially circumferentially spaced from the mode strut, it
may be advantageous to locate the axis 520 relatively closer to the
engine centerline than the axis 524 so as to maintain a good
mechanical advantage for the mode strut.
[0024] An aft end portion 104 of the mode strut is pivotally
coupled to the divergent flap 26 for relative rotation about an
axis 526 fixed relative to the mode strut but floating relative to
the divergent flap with a restricted range of movement. The
exemplary range of movement is provided by the use of a pair of
mounting brackets 110 at an intermediate location on the divergent
flap, each having a slot 112 accommodating an obround slider 113 on
a pivot shaft 114 fixed along the axis 526 relative to the mode
strut. The slider and shaft are free to move along the slot between
first and second ends 116 and 118 thereof. An exemplary
intermediate location is, approximately within the middle third of
the divergent flap length and the middle third of the span between
axes 512 and 522.
[0025] In operation, the position of the synchronization ring 62
determines a nominal throat radius R.sub.T and associated throat
area (i.e., a throttle condition). In a given synchronization ring
position, the aerodynamic forces may then determine the mode which
is nominally associated with the divergent flap interior surface
angle .theta.. FIG. 1 shows the synchronization ring at the forward
extremity of its range of motion, thereby establishing the maximum
nominal throat area. FIG. 1 further shows a high mode condition in
which the aerodynamic forces place the divergent flap in its
maximum .theta. condition with the slider 113 bottomed against the
slot end 116. Under changed conditions, the force balance across
the combination of external flap 50 and divergent flap 26 may
produce an alternate .theta.. For example, in a maximum area,
minimum .theta. low mode condition the slider may be 113 is
substantially bottomed against the slot end 118. In an alternate
configuration, the operation of the mode strut is reversed (i.e.,
the slider arrangement is at the strut's connection to the static
structure rather than at its connection to the divergent flap).
[0026] In minimum throat area/radius conditions, the
synchronization ring 62 is shifted to the rearmost extreme of its
range of motion. During the transition of the synchronization ring,
there is associated telescoping (contraction as shown) of the
external flap. The need to accommodate a sufficient range of
telescoping across the throat area range may, as noted above,
exceed a desired range of extensibility associated with the mode
shift. Thus the mode strut may still operate to restrict a range of
movement of the divergent flap and external flap combination.
[0027] FIG. 1 further shows the upstream flap 22 as including a
main flap 150 and a liner member 152 mounted inboard thereof. The
exemplary liner member 152 includes a panel 154 extending from an
upstream end 156 to a downstream end 158. One or more brackets 160
mount the liner member 152 to the associated main flap 150, holding
the liner member spaced apart therefrom by a gap 162. In operation,
a cooling air flow 164 passes through the gap 162 from an inlet 166
to an outlet 168. The exemplary liner panel 154 has a length that
is a major portion of an overall length of the convergent flap. For
example, the length may be measured as a longitudinal length
L.sub.PB in a max throat high mode or low mode condition wherein
the flap is oriented close to longitudinal. In this condition, the
overall flap length is designated L.sub.FB. Exemplary L.sub.PB is
approximately 80% of L.sub.FB.
[0028] Similarly to the convergent flaps, FIG. 2 shows the
convergent seals 32 as including a main seal member 180 and a liner
member 182. The liner member 182 may have a similar panel and
bracket construction to the liner member 152. Its panel 184 may
have lateral portions configured to interfit and cooperate with
lateral portions of the adjacent panels 154.
[0029] The combined liner members of the convergent flaps and
convergent seals thus forms an overall liner. The liner cooperates
with the convergent flaps and seals to create an interrupted (e.g.,
by the bracket legs) annular channel. The channel carries cool fan
air for discharge into the gas turbine nozzle hot gas stream. The
liner shields the adjacent portions of the convergent flaps and
seals from exhaust gas heating. The discharged air provides a film
cooling effect over the exposed surface of the convergent nozzle
(flaps and seals) and downstream along the throat and divergent
nozzle.
[0030] We have determined that discharging cooling air further
upstream may produce cooling benefits (discussed below). Thus, in
accordance with the present teachings, the liner panel length may
effectively be shortened. FIG. 3 shows a modified convergent flap
222 extending from an upstream end 223 to a downstream end 224. The
convergent flap 222 includes a main flap 226 and a relatively short
liner member 228 at an upstream end thereof. The enlarged view of
FIG. 4 shows the liner member 228 as including a panel 230
extending from an upstream end 232 to a downstream end 234 and
having an inboard surface 236 and an outboard surface 238. The
liner member also includes a mounting bracket 240. The panel has a
longitudinal projected length L.sub.P. An exemplary length L.sub.P
is a minor portion of a projected overall length L.sub.F and a
minor portion of a nearly similar main flap length. In the
exemplary implementation, the main flap length is substantially
close to the unprojected distance/spacing (S) and projected
longitudinal spacing between the axes 502 and 503. The panel 230 is
mounted to an inboard surface 242 of the main flap by the mounting
bracket 240. In the exemplary configuration, the main flap carries
a threaded stud 244 that passes through an aperture in the bracket
240 and is secured to the bracket by a nut 246. A retainer clip 248
mounted to the main flap captures an upstream portion of the
bracket 240 to further retain the liner member. A channel 250 thus
extends between the panel and the main flap from an upstream inlet
252 to a downstream outlet 254 and passes an air flow 256.
[0031] FIG. 5 shows a convergent seal member 260 and a divergent
seal member 262 between adjacent pairs of convergent and divergent
flaps. The convergent seal member 260 includes a main seal 264 and
a similarly foreshortened liner 266 mounted to the main seal 264 in
similar fashion to the mounting of the liner members 228.
[0032] FIG. 6 shows further details of an exemplary liner member
228. The exemplary panel 230 is formed as an assembly of a main
panel element or liner sheet 270, a backing sheet 272, and a flow
blocker 274 secured by rivets 276. The exemplary liner sheet 270
includes a generally flat, rectangular central portion 280 that
extends to the downstream end 224. An arcuate upstream deflector
282 extends upstream from an upstream end of the central portion
280 to the panel upstream end 223. The backing sheet 272 extends
along a forward portion of the central portion 280 and along the
deflector 282, protruding slightly beyond the sides thereof.
Lateral portions 286 and 288 of the liner sheet 270 extend along
lateral edges of the central portion 280 and are curled outboard.
Whereas the exemplary backing sheet 272 lies flat against the liner
sheet 270, the exemplary flow blocker 274 has a main body 290
extending outboard from the liner and backing sheets and a pair of
mounting tabs 292 engaging the outboard surface of the backing
sheet and secured by a lateral two of the rivets 276.
[0033] The exemplary bracket 240 has a central web 300 and lateral
webs or legs 302 and 304 extending outboard. The legs 302 and 304
have upstream end portions 306 extending beyond an upstream end 308
of the web 300 and received in slots 310 in the flow blocker 274
and slots 312 in the backing sheet 272.
[0034] FIG. 6 further shows a recessed area 320 and a central
aperture 322 in the central portion 300 for receiving the mounting
stud and bolt. Additionally, a pair of locating pins 330 (discussed
below) are shown. In an exemplary sequence of manufacture, the
liner sheet 270, backing sheet 272, flow blocker 274, and bracket
240 are made from sheet stock (e.g., by stamping and forming). The
bracket legs are then welded to the outboard surface of the liner
sheet. The welded assembly may then be welded with a protective
coating (e.g., a silicide coating). The backing sheet and flow
blocker are then assembled to the liner sheet with their slots
receiving the bracket legs. The backing sheet and flow blocker may
be pre-coated prior to assembly. Once assembled, the rivets may be
applied.
[0035] For improving alignment of the bracket with the main flap,
the exemplary embodiment utilizes the locating pins 330. These
register with holes 332 in the bracket and corresponding holes (not
shown) in the main flap. The holes 332 may be drilled into the
bracket prior to coating (e.g., after welding). Exemplary materials
for the liner member components are high temperature alloys. In the
exemplary embodiment, the liner sheet 270, bracket 240, rivets 276,
and locating pins 330 are formed of niobium (Nb), the backing sheet
272 is formed of nickel-based superalloy 625, and the flow blocker
274 is formed of nickel-based superalloy 718. The seal liner
members 266 may be similarly manufactured.
[0036] FIGS. 7 and 8 show further details of the assembled liner
member 228. The flow blocker main body 290 has a central recess 336
along its outboard/distal edge 338. The recess 336 may be provided
for clearance relative to the clip 48. The height of the body 290
is selected to provide a desired degree of flow restriction for the
cooling flow 256. The deflector 282 (FIG. 7) has laterally recessed
edge portions 340 beyond which end portions 342 of the backing
sheet protrude. The end portions 342, themselves, have recessed
terminal portions 344 which respectively interfit outboard of
protruding end portions 350 (FIG. 9) of a backing sheet 352 of the
seal liner member 266. FIG. 9 shows the seal liner member 266 as
also including a liner sheet 360 having an upstream deflector 362,
a mounting bracket 364, a flow blocker 366 (FIG. 10) and rivets
368.
[0037] The exemplary liner members 266 include outwardly recessed
lateral portions 370 and 372 defining rebated/recessed areas 374
and 376 shifted outboard of a central portion 378. The recessed
areas 374 and 376 respectively accommodate the lateral portions 288
and 286 to permit interfitting of the respective panels of the
liner members 228 and 266 (FIG. 11). The exemplary recessed areas
374 and 376 are tapered to accommodate the range of convergence of
the nozzle convergent flaps.
[0038] The present nozzle may be engineered as a redesign of an
existing nozzle or otherwise engineered for an existing environment
(e.g., as a drop-in replacement for an existing nozzle such as the
nozzle of FIG. 1 or the convergent flaps and seals thereof). For
example, the illustrated nozzle may be formed as a retrofit kit for
a baseline nozzle such as the nozzle of FIG. 1. In an exemplary
retrofit, the convergent main flaps and main seals and convergent
flap and seal liner members may be replaced. The main flaps and
main seals may be dimensionally similar to the corresponding
baseline components but adapted to include appropriately positioned
mounting studs or other features for the associated liner members.
The liner members may be substantially foreshortened relative to
the corresponding baseline components.
[0039] FIG. 12 shows the engine main (core) exhaust flow 400
meeting the cooling flow 164, 256 at the liner downstream end 158,
234 for the baseline and reengineered nozzles/liners. The cooling
flow provides film cooling along the nozzle surface. The cooling
effectiveness of the film is believed to be a function of: 1) the
distance traveled from the liner (the liner discharge plane); and
2) the mass flux ratio between coolant gas and core gas at the
discharge plane. The coolant gas mixes with the core gas and heats
up as it travels the length of the exposed convergent and divergent
sections. This phenomena is designated as film effectiveness decay,
and is dependent on the mass flux ratio levels at the discharge
plane. As is discussed below, the cut-back liner discharges the
cooling gas film in a more favorable location at a forward portion
of the nozzle. At this forward location the coolant-to-core gas
mass flux ratio is optimized, mixing is reduced, and the enhanced
film effectiveness level offsets the increased length of
travel.
[0040] The effects of this cooling flow may be determined at a
point 402 a distance X along the nozzle downstream of the liner.
The point 402 may be a location of particular criticality (e.g., a
location of maximum temperature or thermal erosion). The point may
be determined experimentally, or simply by post-use observation of
the engine. As the liner is cut back (exit shifted upstream) by a
given distance, the X-value of the particular point will increase
by that distance.
[0041] The flows 400 and 164 each have a density .rho. and a
velocity V. FIG. 13 shows a model plot 412 of the product of this
density and velocity (the mass flux) for the flow 400. FIG. 13 also
shows a model plot 410 of mass flux for the flow 164. The domain
extends from zero to the baseline liner length L.sub.PB. The ratio
of mass flux of the coolant flow 164 to the mass flux of the
exhaust flow 400 is designated the blowing ratio M.
[0042] FIG. 14 shows a model plot 440 wherein the domain is the
ratio of the X-value of the chosen point 402 for the redesigned
(cut-back) liner/nozzle to that value X.sub.B of the baseline
nozzle. The range is the ratio of the blowing ratio M of the
redesigned liner/nozzle to the blowing ratio M.sub.B of the
baseline nozzle. A line 442 divides a first region wherein cooling
at the point 402 is improved (e.g., surface temperature reduced)
from a second region wherein such cooling is reduced. Thus improved
cooling appears to be achieved by increased cut-back.
[0043] The exemplary nozzles have variable throat area (although
the present teachings may also be applied to other nozzles). For a
typical variable nozzle throat area configuration, there is a
partial nesting overlap of lateral portions of the flap liner
panels and seal liner panels. The degree of overlap varies
inversely as a function of nozzle jet area. The interfitting
overlap features (whether actually overlapping in a min. throat
condition, apart in a max. throat condition, or in between) block
flow and induce turbulence, interfering with film cooling
effectiveness. The cut-back may reduce the maximum degree of
overlap and may reduce the extent of the lateral overlap features
thereby reducing or minimizing the amount of film disturbance
generated by the overlap features.
[0044] In the baseline nozzle the cooling air flow is relatively
insensitve to throat condition. This is because the liner exit is
near the throat and the static pressure there is relatively
constant. In the cut-back nozzle, at high nozzle jet areas (e.g.,
at or near the max. throat condition) the core pressure at the
liner exit is reduced relative to an intermediate design throat
area. This reduction causes the fan duct system to flow more
coolant air to the liner system than at the intermediate throat
condition. This enhanced flow rate offsets the enhanced mixing (due
to the higher core velocity air at the high area condition relative
to the intermediate area condition). Similarly, for low nozzle jet
areas, the core velocity is reduced, mixing is reduced, and film
effectiveness enhanced. The enhanced film levels offsets the
reduced flow rate because at low jet areas the liner exit pressure
is increased and less flow is discharged through the liner system.
Therefore the flow and film effectiveness impacts counteract each
other. Thus, as in the baseline, there may be substantial
independence of cooling effectiveness and of nozzle jet area.
[0045] Protection of the convergent flaps/seals, however, imposes
constraints on the cut-back. The cutback exposes a greater portion
of the convergent flaps/seals to exhaust heating. Line 450 of FIG.
14 shows an X/X.sub.B ratio below which (i.e., to the right) the
convergent flap heating exceeds an imposed threshold (e.g., a
maximum temperature greater than a target maximum). In a
redesign/reengineering process, this may determine the chosen liner
length.
[0046] One or more embodiments of the present invention have been
described. Nevertheless, it will be understood that various
modifications may be made without departing from the spirit and
scope of the invention. For example, when implemented as a
reengineering of an existing nozzle, various details of the
existing nozzle may be preserved either by necessity or for
convenience. Additionally, the principles may be applied to
non-axisymmetric nozzles in addition to axisymmetric nozzles and to
vectoring nozzles in addition to non-vectoring nozzles.
Accordingly, other embodiments are within the scope of the
following claims.
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