Turbine engine nozzle
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.
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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.
BACKGROUND OF THE INVENTIONThe invention relates to turbine engines. More particularly, the invention relates to variable throat turbine engine exhaust nozzles.
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 INVENTIONAccordingly, 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.
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
Like reference numbers and designations in the various drawings indicate like elements.
DETAILED DESCRIPTION
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.
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.
In operation, the position of the synchronization ring 62 determines a nominal throat radius RT 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 θ.
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.
Similarly to the convergent flaps,
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.
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.
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.
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.
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 (
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
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.
The flows 400 and 164 each have a density ρ and a velocity V.
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.
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.
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
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.
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 ρv 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.
Type: Application
Filed: Sep 22, 2005
Publication Date: Mar 22, 2007
Applicant:
Inventors: Curtis Cowan (East Hampton, CT), Debora Kehret (Manchester, CT)
Application Number: 11/232,508
International Classification: B63H 11/10 (20060101);