Patent application title: HIGH AREA RATIO TURBINE VANE
Thomas J. Praisner (Colchester, CT, US)
Thomas J. Praisner (Colchester, CT, US)
Eunice Allen-Bradley (Vernon, CT, US)
Norbert Huebner (Dachau, DE)
Norbert Huebner (Dachau, DE)
IPC8 Class: AF01D514FI
Class name: Fluid reaction surfaces (i.e., impellers) specific blade structure (e.g., shape, material, etc.) turbo machine
Publication date: 2012-11-01
Patent application number: 20120275922
A vane for a turbine engine comprises an airfoil section, an inner
platform and an outer platform. The airfoil section comprises pressure
and suction surfaces extending from a leading edge to a trailing edge.
The inner platform is attached to the airfoil section along an inner flow
boundary, where the inner flow boundary extends from an upstream inlet
region of the vane to a downstream outlet region of the vane. The outer
platform is attached to the airfoil section along an outer flow boundary,
where the outer flow boundary extends from the inlet region to the outlet
region. An area ratio of the outlet region to the inlet region is greater
1. A vane for a turbine engine, the vane comprising: an airfoil section
comprising pressure and suction surfaces extending from a leading edge to
a trailing edge; an inner platform attached to the airfoil section along
an inner flow boundary, the inner flow boundary extending from an
upstream inlet region of the vane to a downstream outlet region of the
vane; and an outer platform attached to the airfoil section along a outer
flow boundary, the outer flow boundary extending from the upstream inlet
region to the downstream outlet region; wherein an area ratio of the
outlet region to the inlet region is greater than 2.4.
2. The vane of claim 1, wherein the area ratio comprises a ratio of a cross-sectional area of the outlet region between the inner flow boundary and the outer flow boundary, to a cross-sectional area of the inlet region between the inner flow boundary and the outer flow boundary.
3. The vane of claim 2, wherein the area ratio is between 2.4 and 2.6.
4. The vane of claim 2, wherein the area ratio is between 2.5 and 2.9.
5. The vane of claim 1, wherein an aspect ratio of the vane is less than 2.0.
6. The vane of claim 5, wherein the aspect ratio is 1.5 or less.
7. The vane of claim 6, wherein a radial spacing between the inner flow boundary and the outer flow boundary varies between the inlet region and the outlet region.
8. A turbine comprising the vane of claim 1, wherein the vane defines a portion of a flow duct about a turbine axis.
9. The turbine of claim 8, wherein the airfoil section defines a local flow angle of at least ten degrees with respect to the turbine axis.
10. A turbine vane comprising: an airfoil section comprising a leading edge, a trailing edge, and pressure and suction surfaces defined therebetween; an inner platform attached to the airfoil section, the inner platform defining an inner flow boundary between an inlet region of the turbine vane and an outlet region of the turbine vane; and an outer platform attached to the airfoil section, the outer platform defining an outer flow boundary between the inlet region of the turbine vane and the outlet region of the turbine vane; wherein a ratio of a cross sectional area of the outlet region to a cross sectional area of the inlet region is greater than 2.4.
11. The turbine vane of claim 10, wherein the ratio is less than 2.9
12. The turbine vane of claim 11, wherein the ratio is at least 2.5.
13. The turbine vane of claim 10, wherein an aspect ratio of the airfoil section is 1.5 or less.
14. A turbine engine comprising a plurality of turbine vanes as specified in claim 10, the plurality of turbine vanes defining an annular flow duct about a turbine axis.
15. The turbine engine of claim 14, wherein the plurality of turbine vanes define a duct angle of at least ten degrees with respect to the turbine axis.
16. A turbine engine comprising: a plurality of vanes arranged circumferentially about a turbine axis, each of the vanes comprising an airfoil, an inner flow boundary defined along an inner radius of the airfoil, and an outer flow boundary defined along an outer radius of the airfoil; and an annular flow duct defined between the inner flow boundary and the outer flow boundary, the annular flow duct having an inlet region and an outlet region; wherein a ratio of a cross sectional area of the outlet region to a cross sectional area of the inlet region is between 2.4 and 2.9.
17. The turbine engine of claim 16, wherein each of the airfoils has an aspect ratio of 1.5 or less.
18. The turbine engine of claim 17, wherein the annular flow duct has a duct angle of at least ten degrees with respect to the turbine axis.
19. The turbine engine of claim 18, further comprising: a first turbine section upstream of the turbine section; and a second turbine section downstream of the turbine section; wherein the annular flow duct comprises a transition duct between the first turbine section and the second turbine section.
20. A turbofan engine comprising the turbine engine of claim 19 in flow series with a compressor and a propulsion fan.
 This invention relates generally to turbine engines, and specifically to turbine vane design. In particular, the invention concerns airfoil geometry for turbine vanes.
 Gas turbine engines provide reliable, efficient sources of power and mechanical energy for use in aviation, electrical power generation, industrial heating and cooling, hydrocarbon fuel production, naval and marine vessels, and a range of other fluid processing needs. Energy is produced by combustion in a turbine core comprising a compressor, combustor and turbine in flow series with an upstream inlet and a downstream exhaust.
 The compressor compresses air from the inlet. The compressed air is mixed with fuel in the combustor, and ignited to produce hot combustion gases. The combustion gases expand in the turbine, which drives the compressor via a common shaft, and then exit via the exhaust. Power is delivered in the form of rotational energy in the shaft or reactive thrust from the exhaust, or both.
 In many gas turbines, a number of different turbine and compressor sections are arranged into a series of coaxially nested spools, which operate at different pressures and rotational speeds. The compressor and turbine sections are subdivided into a series of stages, which are formed by alternating rows of rotor blade and stator vanes. The blades and vanes include airfoil sections that are shaped to compress, turn and accelerate incoming air in the compressor section, and to generate lift and extract rotational energy from expanding combustion gas in the turbine section.
 Typical compressor and turbine designs include axial, centrifugal and axial/radial flow stages, with both direct and indirect shaft couplings. Depending on configuration, the compressor may also provide cooling air for downstream engine components, including high-pressure compressor and turbine airfoils, and other elements of the combustor, turbine and exhaust that are exposed to hot working fluid flow.
 Ground-based industrial gas turbines (IGTs) can be quite large, and may utilize complex spooling systems for increased efficiency. Some industrial turbines are configured for combined-cycle operation, in which additional energy is extracted from the partially-cooled exhaust gas stream, for example in a steam turbine. Industrial gas turbines deliver power via an output shaft connected to a mechanical load such as an electrical generator, blower or pump.
 Aviation applications include turbojet, turbofan, turboprop and turboshaft engines. Turbojet engines are an older design, in which thrust is generated primarily from the exhaust. Modern fixed-wing aircraft typically employ turbofan and turboprop configurations, in which the low spool is coupled to a propulsion fan or propeller.
 In turboprop aircraft, a reduction gearbox is used to reduce tip speed. For turboshaft engines, which are used on rotary wing aircraft, the reduction ratio is generally higher. Advanced turbofan engines may also include geared drive mechanisms, providing independent fan speed control for reduced engine noise and increased operating efficiency.
 Most commercial jets are powered by two- or three-spool high-bypass turbofan engines. High-bypass turbofans generate most of their thrust from the fan, which drives bypass flow through a duct oriented around the engine core. Unducted designs are also known, including counter-rotating and aft-mounted fan configurations.
 Low-bypass turbofans produce higher specific thrust, but at some cost in noise and fuel efficiency. Low-bypass turbofans are used on supersonic military jets and other high-performance aircraft, and are commonly configured for thrust augmentation or afterburning. Generally, afterburners are limited to short-duration usage because of the increased operational stress and high rate of fuel consumption, but continuously afterburning turbojet designs are known.
 In aviation applications, the main engines also provide power for accessory functions including pneumatics, hydraulics and environmental control, either via the bleed air system or by driving an accessory gearbox coupled to an electrical generator (or both). Alternatively, an auxiliary power unit (APU) is used. APU systems are based on small-scale (usually one-spool) turbine cores, which operate to generate electrical power and supply cabin air while the aircraft is on the ground. APUs can also provide compressed air for main engine startup, and may be configured for in-flight operation to provide an independent or emergency backup power source for hydraulics, avionics and flight control systems.
 Across these applications, turbine performance depends on precise control of the working fluid flow, including a detailed understanding of the various subsonic, transonic and supersonic flow components along each section, between individual airfoils, and along adjacent flow boundaries. In particular, lift, acceleration and turning efficiency depend upon on a range of non-linear effects, including turbulence, laminar flow separation, laminar/turbulent boundary transitions, vorticity, shock wave formation and shock interference.
 As a result, turbine blade, vane and airfoil design is a highly complex, specialized and unpredictable art, requiring continuous tradeoffs between a range of difficult to quantify flow effects and associated loss mechanisms. Actual part design is further complicated by cooling flow requirements and external constraints on size, weight and material composition, as compared to the required thrust output, and by additional reliability and durability concerns over a wide range of working fluid pressures and temperatures, flow rates, and engine spool speeds.
 This invention concerns a vane for a gas turbine engine. The vane includes an airfoil section with pressure and suction surfaces extending between a leading edge and a trailing edge. An inner platform is attached to the airfoil section along the inner flow boundary, and an outer platform is attached along the outer flow boundary.
 The inner and outer flow boundaries extend from an upstream inlet region of the vane to a downstream outlet region of the vane. The area ratio of the outlet region to the inlet region is greater than 2.4.
BRIEF DESCRIPTION OF THE DRAWINGS
 FIG. 1 is a cross-sectional view of a gas turbine engine with a high area ratio vane ring.
 FIG. 2 is an axial view of the vane ring.
 FIG. 3 is a schematic side view of a vane for the vane ring.
 FIG. 1 is a cross-sectional view of gas turbine engine 10, in a turbofan embodiment. In this embodiment, turbine engine 10 comprises fan 12, bypass duct 14, compressor section 16, combustor 18 and turbine section 20. Flow FI from inlet 22 is accelerated by fan 12 to generate bypass flow FB through bypass duct 14 and core flow FC through compressor section 16, combustor 18 and turbine section 20. Exit flow FE is exhausted from turbine engine 10 at outlet (exhaust nozzle) 24.
 For a typical two-spool configuration, compressor section 16 comprises low pressure compressor (LPC) 26 and high pressure compressor (HPC) 28. Turbine 20 comprises high pressure turbine (HPT) 30 and low pressure turbine (LPT) 32. Compressor 16 and turbine 20 are further divided into a number of stages, which are formed by alternating rows of stator vanes 34 and rotor blades 36. Each vane 34 and blade 36 includes an airfoil section with geometry selected for low loss, high efficiency acceleration, turning, lift, compression and expansion of the working fluid flow.
 In at least one stage of compressor 16 or turbine 20, vanes 34 form a flow duct with a high area ratio for improved engine performance with reduced flow separation, as described below. In some embodiments, high area ratio vanes 34 are arranged into a turbine vane ring, a transition duct or a turning mid-turbine frame (TMTF) structure positioned between first and second turbine sections, for example high pressure turbine section 30 and low pressure turbine section 32.
 In the two-spool, high-bypass configuration of FIG. 1, high pressure shaft 38 couples high pressure compressor 28 to high pressure turbine 30, forming the high pressure (HP) spool or high spool. Low pressure shaft 40 couples low pressure compressor 26 and fan 12 to low pressure turbine 32, forming the low pressure (LP) spool or low spool. The high and low spools are coaxially mounted about turbine axis (centerline) CL, with high pressure shaft 38 and low pressure shaft 40 rotating at different speeds.
 Fan 12 comprises a number of fan blades arranged around a disk or other rotating member. Depending on embodiment, fan 12 is configurable for high-bypass operation in a commercial or regional jet, or for low-bypass operation in a high-performance aircraft such as a military jet fighter. In some engines, fan 12 is coupled to the low spool (low pressure shaft 36) via geared fan drive mechanism 42, providing additional fan speed control. Alternatively, fan 12 is directly coupled to low pressure compressor 26, and co-rotates with low pressure shaft 38.
 In operation of turbine engine 10, high pressure turbine 30 drives high pressure compressor 28. Low pressure turbine 32 drives low pressure compressor 26 and fan 12, which generates thrust by accelerating incoming airflow F1 from inlet 22 to create bypass flow FB through bypass duct 14. Low pressure compressor 26 and high pressure compressor 28 compress core flow FC for combustor 18.
 Fuel is injected into combustor 18, where it mixes with the compressed air and is ignited to produce hot combustion gas. The hot combustion gas exits combustor 18 to enter high pressure turbine 30 and low pressure turbine 32, which generate rotational energy from the expanding combustion gas. Exit flow FE generates additional thrust at exhaust nozzle 24.
 Depending on configuration, low pressure compressor 32 may be omitted, or may function as an intermediate pressure (IP) compressor. Alternatively, turbine engine 10 utilizes a three-spool design with separate low, intermediate and high pressure spools. In other embodiments, turbine engine 10 comprises an unducted turbofan, turboprop or turboshaft engine, an auxiliary power unit, or an industrial gas turbine, as described above, and the spool and shaft configurations vary accordingly.
 Across these different embodiments, thermodynamic efficiency depends on overall pressure ratio; that is, the ratio of stagnation pressure downstream of compressor 16 and upstream of fan 12. In turbine section 20, performance also depends on the area ratio, which is the ratio of cross-sectional flow areas at the outlet and inlet of the flow duct formed by turbine vanes 34. In particular, expanding the flow path to larger area is beneficial for downstream turbine efficiency, and reduces vibratory stress on rotor blades 36 and other turbine components.
 FIG. 2 is an axial view of turbine vane ring 50, taken in a downstream direction along centerline CL of FIG. 1. Turbine vane ring 50 comprises a plurality of vanes 34, arranged circumferentially about turbine axis CL.
 Each vane 34 extends along an airfoil section from inner platform 52 at inner diameter (ID) flow boundary 54 to outer platform (or shroud) 56 at outer diameter (OD) flow boundary 58. The airfoil sections are formed by convex (suction) and concave (pressure) surfaces, which extend between leading and trailing edges defined by the working fluid flow.
 In one embodiment, individual turbine vanes 34 are attached to a static turbine support structure to form ID flow boundary (or ID flow ring) 54 along adjacent inner platforms 52, and OD flow boundary (or OD flow ring) 58 along adjacent outer platforms 56. Alternatively, vanes 34 are assembled or manufactured in pairs, triplets or other groupings, and turbine ring 50 may be formed in axially split ring sections, or as a complete annular structure. In some of these configurations, one or both of inner platforms 52 and outer platforms 56 are formed as part ID and OD flow boundaries 56 and 58, respectively, and the airfoil sections are separate components.
 In turbine embodiments, vanes 34 may be formed of high-temperature materials such as nickel, cobalt and iron-based alloys, super alloys or ceramic matrix composites. In addition, a bond coat, thermal barrier coating (TBC) or other protective layer may be applied to surfaces exposed to high-temperature working fluid flow.
 In lower-temperature applications, for example compressor sections, vanes 34 may be formed of a strong, lightweight metal such as titanium or aluminum, or an alloy thereof. Alternatively, vanes 34 comprise other metals and metal alloys, or graphite, polymers and other composite materials.
 In high-temperature applications, vanes 34 are provided with interior passages for impingement or film cooling, or to transport cooling fluid from one turbine component to another. Alternatively, vanes 34 are sized to accommodate additional turbine hardware including tie rods, struts, bearing supports, oil lines and electrical wires for sensors, actuators and controllers.
 FIG. 3 is a schematic side view of turbine vane 34, in circumferential projection about axial centerline CL of turbine vane ring 50, as shown in FIG. 2. Vane 34 comprises inner platform 52 at ID boundary 54, outer platform 56 at OD boundary 58, and airfoil section 60.
 Airfoil portion (or airfoil) 60 extends in a spanwise sense from inner platform 52 to outer platform 56, and axially from leading edge 62 to trailing edge 64. Leading edge 62 and trailing edge 64 define leading edge (LE) span S1 and trailing edge (TE) span S2, respectively.
 Inner platform 52 and outer platform 56 extend along ID boundary 54 and OD boundary 58 between inlet region 66 and outlet region 68, providing an improved flow transition between upstream rotor stage 36A and downstream rotor stage 36B. In particular, turbine vane 34 provides a high area ratio between downstream outlet 68 and upstream inlet 66, expanding the flow path to higher annular area. Vane 34 reduces flow detachment, flow separation, turbulence, voracity and other loss and vibration-inducing effects, including vibratory effects on rotor stage 36B, immediately downstream of vane 34.
 In the particular embodiment of FIG. 3, turbine vane 34 comprises a turning mid-turbine frame vane positioned in a transition duct between a high-pressure turbine section with immediately upstream rotor blade 36A, and a low-pressure turbine section with immediately downstream rotor blade 36B. Alternatively, vane 34 is positioned between high, low, or intermediate sections of a turbine or compressor assembly, or between two rotor stages within a particular turbine or compressor section.
 In each of these embodiments, airfoil section 60, inner platform 52 and outer platform 56 of vane 34 are configured for efficient turning, acceleration (or deceleration) and expansion (or compression) of the working fluid flow, between inner flow boundary 54 and outer flow boundary 58. Stator vane 34 also encompasses aerodynamic fairing embodiments, in which turning and acceleration are reduced, but the advantages of increased flow area without substantial flow separation are retained.
 In the circumferential projection of FIG. 3, leading edge span S1 and trailing edge span S2 include radial, axial and circumferential components, accounting for forward or aft axial sweep and tangential lean or curvature. Chord length C is defined between mid-span point M1 of leading edge 62, and mid-span point M2 of trailing edge 64.
 Duct angle θ is the angle between chord C and the direction of turbine axis CL, and axial chord length L is the axial projection of chord C:
L=C×cos θ. 
 The aspect ratio (AS) of airfoil section 60 is defined by the average of leading edge span S1 and trailing edge span S2, divided by axial chord length L:
AS = ( S 1 + S 2 ) 2 L . [ 2 ] ##EQU00001##
The area ratio (AR) of vane 34 (and turbine vane ring 50, see FIG. 2) is the ratio of annular flow area A2 at outlet 68 to annular flow area A1 at inlet 66. Specifically:
AR = A 2 A 1 . [ 3 ] ##EQU00002##
 Inlet and outlet areas A1 and A2 are defined by the cross-sectional flow area between ID flow boundary 54 and OD flow boundary 58, respectively. For annular flow duct configurations, this is: and
Inner flow duct radius RID is the radial distance from turbine axis CL to ID flow boundary 54, and outer radius ROD is the distance from turbine axis CL to OD flow boundary 58.
 For inlet area A1, inner radius RID is defined along ID flow boundary 54 at the upstream (axially forward) end of inner platform 52, and outer radius ROD is defined along OD flow boundary 58 at the upstream (axially forward) end of outer platform 56. Alternatively, inlet area A1 is defined between ID flow boundary 54 and OD flow boundary 58 in the plane of mid-span point M3 on the trailing edge of inlet rotor blade 36A, in the rotor stage immediately upstream of vane 34.
 Similarly, outlet area A2 is defined between ID flow boundary 54 and OD flow boundary 58 at the downstream (axially aft) ends of inner platform 52 and outer platform 56, respectively. Alternatively, outlet area A2 is defined in the plane of mid-span point M4 on the leading edge of outlet rotor blade 36B, in the first rotor stage immediately downstream of airfoil 60. A particular area ratio AR can be achieved with a number of uniform vanes 34 all having substantially the same geometry, including aspect ratio AS of airfoil 60, or a number of vanes 34 having different airfoil sections 60 with different aspect ratios AS.
 For positive duct angles θ, chord C diverges or extends away from turbine axis CL in the downstream axial direction, and area ratio AR is generally larger than one because the duct area goes as the square of the radius. Area ratio AR also increases with overall axial length D of vane 34, where axial length D includes both axial chord length L of airfoil 60 and the upstream and downstream contributions of inner platform 52 and outer platform 56.
 Larger area ratios AR represent increased radial height along the downstream flow path, reducing endwall losses that account for turbine inefficiency. Vane 34 also provides a more favorable or optimal flow path between the upstream and downstream rotor stages, as described above, reducing flow separation and improving performance while reducing vibration in downstream rotor stage 36B.
 Specifically, turbine vane 34 provides a relatively high area ratio AR of greater than 2.4, up to about 2.9. In some embodiments, area ratio AR is about 2.5, or between 2.5 and 2.9, and in other embodiments area ratio AR is at least 2.7, for example between 2.7 and 2.9. In further embodiments, area ratio AR is greater than 2.9, for example about 3.0, or higher.
 Area ratio AR increases with (positive) duct angle θ. In some embodiments, duct angle (or local flow path angle) 0 is about ten degrees (10°). Alternatively, duct angle θ is at least ten degrees, for example between ten degrees and fifteen degrees (10-15°), or between ten degrees and twenty degrees (10-20°). In additional embodiments, duct angle θ is greater than five degrees (5°), for example between five and fifteen degrees (5-15°), or between five and twenty degrees (5-20°), or between ten and thirty degrees (10-30°). Alternatively, duct angle θ is larger, for example 25-35° or 20-40°, or 30-45°.
 For positive duct angles θ, area ratio AR also increases with chord length C (and axial chord length L). For given span heights S1 and S2, this corresponds to a reduced aspect ratio AS.
 In some embodiments, vane 34 comprises a long chord airfoil 60, with aspect ratio AS of about 1.5, or between about 1.4 and about 1.6. In other embodiments, aspect ratio AS is less than about 1.5 or 1.6, for example between 1.2 and 1.6, or between 1.0 and 1.5. In further embodiments, aspect ratio AS is less than 2.0, for example between 1.0 and 2.0.
 Area ratio AR also depends on the relative divergence or convergence of ID flow boundary 54 and OD flow boundary 58. Relative convergence and divergence are defined by the increase or decrease in radial spacing ROD-RID, as defined between inlet region 66 and outlet region 68.
 In the diverging case, where radial spacing ROD-RID (between ID boundary 54 and OD boundary 58) increases from inlet region 66 to outlet region 68, the effect on area ratio AR is also increased. Because of the r-squared effect of duct radius, however, area ratio AR also increases with axial length D when ID and OD boundaries 54 and 58 are substantially parallel; that is, with substantially constant radial spacing (for example, within five or ten percent, including small variations in either direction along the flow duct). In fact, area ratio AR may also increase when ID and OD boundaries 54 and 58 converge between inlet 66 and outlet 68 (that is, with decreasing radial spacing).
 These aspects of high area ratio vane 34 are the result of an aggressive design process, requiring a number of complex tradeoffs to achieve precise flow control and maintain flow diffusion without substantial (or even massive) flow separation. In particular, the structure of vane 34 is defined by design choices involving the effects of turbulence, laminar flow transitions, vorticity and shock wave formation, for which the corresponding loss mechanisms are complex, non-linear, and difficult to accurately model or predict.
 While this invention has been described with reference to exemplary embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the spirit and scope of the invention. In addition, modifications may be made to adapt a particular situation or material to the teachings of the invention, without departing from the essential scope thereof. Therefore, the invention is not limited to the particular embodiments disclosed herein, but includes all embodiments falling within the scope of the appended claims.
Patent applications by Eunice Allen-Bradley, Vernon, CT US
Patent applications by Norbert Huebner, Dachau DE
Patent applications by Thomas J. Praisner, Colchester, CT US
Patent applications in class Turbo machine
Patent applications in all subclasses Turbo machine