Patent Application
20110070075
This invention relates generally to heat transfer in gas turbine engines and more particularly to apparatus for cooling structures in such engines.
A gas turbine engine includes a turbomachinery core having a high pressure compressor, combustor, and high pressure turbine (“HPT”) in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. The high pressure turbine includes annular arrays (“rows”) of stationary vanes or nozzles that direct the gases exiting the combustor into rotating blades or buckets. Collectively one row of nozzles and one row of blades make up a “stage”. Typically two or more stages are used in serial flow relationship. The combustor and HPT components operate in an extremely high temperature environment, and must be cooled by air flow to ensure adequate service life.
Cooling air flow is typically provided by utilizing relatively lower-temperature “bleed” air extracted from an upstream part of the engine, for example the high pressure compressor, and then feeding that bleed air to high-temperature downstream components. The bleed air may be applied in numerous ways, for example through internal convection cooling or through film cooling. When used for convection cooling, the bleed air is often routed through serpentine passages or other structures which generate a pressure loss as the cooling air passes through them. Because bleed air represents a loss to the engine cycle and reduces efficiency, it is desired to maximize heat transfer rates and thereby use the minimum amount of cooling flow possible. For this reason heat transfer improvement structures, such as turbulence promoters or “turbulators”, are often formed on cooled surfaces.
Turbulators are elongated strips or ribs having a square, rectangular, or other symmetric cross-section, and are generally aligned transverse to the direction of flow. The turbulators serve to “trip” the boundary layer at the component surface and create turbulence which increases heat transfer. Cooling effectiveness is thereby increased. One problem with the use of conventional turbulators is that a flow stagnation zone is present downstream of each turbulator. This zone causes dust, which is naturally entrained in the cooling air, to be deposited and build up behind the turbulators. This build-up is an insulating layer which reduces heat transfer also can cause undesirable wear.
An example of a particular gas turbine engine structure requiring effective cooling is an HPT nozzle. HPT nozzles are often configured as an array of airfoil-shaped vanes extending between annular inner and outer bands which define the primary flowpath through the nozzle. Some prior art HPT nozzles have experienced temperatures on the aft inner band above the design intent. This has lead to the loss of the aft inner band because of oxidation at a low number of engine cycles. The material loss can trigger a chain of undesirable events, leading to serious engine failures. For example, in a multi-stage HPT, the loss of the aft portion of the first stage nozzle inner band can cause hot gas ingestion between the first stage nozzle and the forward rotating seal member or “angel wing” of the adjacent first stage blade. The ingested primary flow can in turn heat up the forward cooling plate of the first stage rotor disk causing it to crack. Once the cooling plate is cracked, hot air can heat up the first stage rotor disk causing damage to the disk post, which could lead to the release of a first stage turbine blade.
These and other shortcomings of the prior art are addressed by the present invention, which provides a “fastback” turbulator structure that discourages stagnation of high velocity flow.
According to one aspect of the invention, a heat transfer apparatus includes: (a) a member defining a wall exposed to fluid flow in a predetermined direction of flow; and (b) a plurality of turbulators disposed on the wall, each turbulator having: (i) an upright front face which generally faces the direction of flow, and (ii) a back face which defines a ramp-like shape tapering from the front face to the wall.
According to another aspect of the invention, a turbine nozzle includes: (a) a hollow, airfoil-shaped turbine vane; (b) an arcuate first band disposed at a first end of the turbine vane, the first band having a flowpath face adjacent the turbine vane, and an opposed back face; (c) wherein the back face includes at least one open pocket, the at least one pocket defined in part by a bottom wall recessed from the back face, opposed ends of the bottom wall merging with the back face, where the pocket is exposed to fluid flow in a predetermined direction of flow; and (d) a plurality of turbulators disposed on the bottom wall, each turbulator having: (i) an upright front face which generally faces the direction of flow, and (ii) a back face which defines a ramp-like shape tapering from the front face to the bottom wall of the pocket.
The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which:
Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,
In the illustrated example, the engine is a turbofan engine and a low pressure turbine (not shown) would be located downstream of the gas generator turbine 10 and coupled to a shaft driving a fan. However, the principles described herein are equally applicable to turboprop and turbojet engines, as well as turbine engines used for other vehicles or in stationary applications.
The high pressure turbine 10 includes a first stage nozzle 12 which comprises a plurality of circumferentially spaced airfoil-shaped hollow first stage vanes 14 that are supported between an arcuate, segmented first stage outer band 16 and an arcuate, segmented first stage inner band 18. The first stage vanes 14, first stage outer band 16, and first stage inner band 18 are arranged into a plurality of circumferentially adjoining nozzle segments that collectively form a complete 360° assembly. The first stage outer and inner bands 16 and 18 define the outer and inner radial flowpath boundaries, respectively, for the hot gas stream flowing through the first stage nozzle 12. The first stage vanes 14 are configured so as to optimally direct the combustion gases to a first stage rotor 20.
The first stage rotor 20 includes an array of airfoil-shaped first stage turbine blades 22 extending outwardly from a first stage disk 24 that rotates about the centerline axis of the engine. A segmented, arcuate first stage shroud 26 is arranged so as to closely surround the first stage turbine blades 22 and thereby define the outer radial flowpath boundary for the hot gas stream flowing through the first stage rotor 20.
A second stage nozzle 28 is positioned downstream of the first stage rotor 20, and comprises a plurality of circumferentially spaced airfoil-shaped hollow second stage vanes 30 that are supported between an arcuate, segmented second stage outer band 32 and an arcuate, segmented second stage inner band 34. The second stage vanes 30, second stage outer band 32 and second stage inner band 34 are arranged into a plurality of circumferentially adjoining nozzle segments that collectively form a complete 360° assembly. The second stage outer and inner bands 32 and 34 define the outer and inner radial flowpath boundaries, respectively, for the hot gas stream flowing through the second stage turbine nozzle 34. The second stage vanes 30 are configured so as to optimally direct the combustion gases to a second stage rotor 38.
The second stage rotor 38 includes a radial array of airfoil-shaped second stage turbine blades 40 extending radially outwardly from a second stage disk 42 that rotates about the centerline axis of the engine. A segmented arcuate second stage shroud 44 is arranged so as to closely surround the second stage turbine blades 40 and thereby define the outer radial flowpath boundary for the hot gas stream flowing through the second stage rotor 38.
The inner band 18 has a flowpath face 54 and an opposed back face 56. One or more open pockets 58 are formed in the back face 56. The pockets 58 may be formed by incorporating them into the casting, by machining, or by a combination of techniques.
The bottom wall 66 extends in a generally circumferential direction between first and second ends 68 and 70. The bottom wall 66 includes a central portion 72 which is recessed from the back face 56 and two end portions 74. The end portions 74 form ramps between the central portion 72 and the back face 56. The central portion 72 may define a portion of a circular arc, or another suitable curved profile.
The distance that the bottom wall 66 is offset from the back face 56 in a radial direction is referred to as the “depth” of the pocket 58 and is denoted “D”. The specific value of “D” varies at each location of the pocket 58, generally being the greatest near the circumferential midpoint of the pocket 58 and tapering to zero at the ends 68 and 70. It is desirable for weight reduction purposes to make the depth “D” as large as possible. The maximum depth achievable is limited by the minimum acceptable material thickness in the inner band 18 and the vane 14, shown at “T” (see
As shown in
Unlike prior art turbulators described above, the turbulators 100 are shaped so as to avoid flow stagnation and dust buildup. In particular, with reference to
The peak height “H” of the turbulator 100 above the bottom wall 66 is selected in accordance with prior art practice, and is large enough so that each turbulator 100 is effective in producing turbulence, that is, the turbulator 100 is significantly taller than surface imperfections in the cast component surface, but generally not so large as to form a significant flow blockage. For example, the height “H” may be from about 0.18 mm (0.007 in.) to about 0.64 mm (0.025 in.). A height of about 0.25 mm (0.010 in.) is believed to be a preferred value in the specific example illustrated.
The turbulators 100 are spaced-apart from each other in the direction of cooling air flow by a distance “S” which is selected to suit the specific application. As a general rule of thumb, the distance S may be about 8 to 10 times the height H.
As illustrated, the back face 104 is substantially planar over the majority of its surface and is inclined so as to give the turbulator 100 an included angle φ. The angle φ is selected to be large enough so that each turbulator 100 has a reasonable overall length (i.e. in the direction of cooling air flow), but not so large that a stagnation zone would be present during operation. As an example, the angle φ may be about 20° or less. It is believed that an angle φ of about 7° is a preferred value for preventing recirculation. The back face 104 of each turbulator 100 may extend all the way to the root of the front face 102 of the downstream turbulator, or it may terminate at a shorter distance, leaving an exposed portion of the bottom wall 66 between each turbulator 100.
The turbulator 100 need not have a planar shape; for example the back face could be curved in a convex, airfoil-like shape (not shown) so as to maximize Coanda effect in the flow over the turbulator 100 and further discourage flow separation.
In operation, a substantial purge flow of relatively cool air occurs in the secondary air flow path in contact with the back face 56 of the inner band 18. Its velocity is primarily tangential (i.e. into or out of the page in
The “fastback” turbulator structures described above are useable not only in turbine nozzles, but also for any structure requiring heat transfer enhancement, in particular in any structure where prior art turbulators might otherwise be used. Nonlimiting examples of such structures include gas turbine engine combustor liners, stationary (i.e. frame) structures, turbine shrouds and hangers, turbine disks and seals, and the interiors of stationary or rotating engine airfoils such as nozzles and blades. As such, the components described above should be considered as merely one example representative of a heat transfer structure having a wall exposed to fluid flow with turbulators disposed thereon. The fastback turbulators may be incorporated into the casting of a component, may be machined into an existing surface, or may be provided as separate structures which are then attached to a surface. They are believed to be particularly effective in regions of high-speed flow, and where swirl flow dominates.
The foregoing has described a fastback turbulator structure and a pocket geometry for a turbine nozzle band. While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the spirit and scope of the invention. Accordingly, the foregoing description of the preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of illustration only and not for the purpose of limitation.
This application claims the benefit of Provisional Patent Application 61/245,649, filed Sep. 24, 2009.
| Number | Date | Country | |
|---|---|---|---|
| 61245649 | Sep 2009 | US |
