Patent Application
20180017019
This invention relates generally to gas turbine engines and more particularly to the fans of such engines.
A gas turbine engine includes, in serial flow communication, a compressor, a combustor, and turbine. The turbine is mechanically coupled to the compressor and the three components define a turbomachinery core. The core is operable in a known manner to generate a flow of hot, pressurized combustion gases to operate the engine as well as perform useful work such as providing propulsive thrust or mechanical work. A turbofan engine includes, in addition to the core, a low-pressure turbine coupled to a fan configured to produce a bypass stream.
It is desirable to improve turbofan engine performance over wider operating ranges than is presently possible. One way this can be achieved is by using an auxiliary convertible fan stage to enable variations in fan bypass and pressure ratio levels. However, this apparatus is relatively complex and heavy and requires additional moving parts in the fan module.
An alternate form of inducing similar variations in fan bypass and pressure ratio levels is to “overflow” the fan forcing a reduction in fan operating line and pressure ratio level. A side effect of an overflow condition is a fan efficiency decrease. This side effect can be mitigated by incorporating low solidity levels in the rear stage rotor and stator, providing aerodynamic choking relief.
One problem with a fan having low solidity is that the rotor hub and stator endwalls are susceptible to flow separation.
This problem is addressed by a turbofan engine incorporating splitter airfoils into the fan.
According to one aspect of the invention, A turbofan engine includes: a turbomachinery core operable to produce a flow of combustion gases; a low-pressure turbine configured to extract energy from the combustion gases so as to drive a fan to produce a fan flow, the fan being configured such that at least a portion of the fan flow exits the engine without passing through a turbine; wherein the fan includes: a rotor having at least one rotor stage including a rotatable disk defining a rotor flowpath surface and an array of axial-flow rotor airfoils extending outward from the flowpath surface; at least one stator stage having a wall defining a stator flowpath surface, and an array of axial-flow stator airfoils extending away from the stator flowpath surface; and wherein at least one of the rotor or stator stages includes an array of airfoil-shaped splitter airfoils extending from at least one of the flowpath surfaces thereof, the splitter airfoils alternating with the rotor or stator airfoils of the corresponding stage, wherein at least one of a chord dimension of the splitter airfoils and a span dimension of the splitter airfoils is less than the corresponding dimension of the airfoils of the at least one stage.
According to another aspect of the invention, a method of operating a variable-cycle gas turbine engine includes: using a turbomachinery core including in sequential flow relationship: a compressor, a combustor, and a turbine mechanically coupled to the compressor to generate a flow of combustion gases; using a low-pressure turbine to extract energy from the combustion gases so as to drive a fan to produce a fan flow, the fan being configured such that at least a portion of the fan flow exits the engine without passing through a turbine, wherein the fan incorporates at least one row of splitter airfoils; and during engine operation, using at least one variable-cycle device to vary a backpressure downstream of the fan, thereby moving an operating line of the fan by at least 5% from a nominal position.
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,
An outer casing 36 is spaced apart from the core 30 by an inner annular wall 38 so as to define an annular bypass duct 40 therebetween. The outer casing 36 defines an inlet 42 at its upstream end.
The fan 12 may include a number of rotor stages, each of which comprises a row of fan blades 44 mounted to a rotor 46. The fan 12 also includes at least one stator stage comprising a row of stationary airfoils that serve to turn the airflow passing therethrough. In the illustrated example the fan 12 includes inlet guide vanes 48 upstream of the rotor 46, stator vanes 50 disposed between rotor stages, and outlet guide vanes 52 downstream of the rotor 46. For purposes of the present application, any of the inlet guide vanes 48, the stator vanes 50, and the outlet guide vanes 52 may be considered to be “stator airfoils”. The fan 12 is configured for axial fluid flow, that is, fluid flow generally parallel to the centerline axis 11. This is in contrast to a centrifugal compressor or mixed-flow compressor.
It is noted that, as used herein, the terms “axial” and “longitudinal” both refer to a direction parallel to the centerline axis 11, while “radial” refers to a direction perpendicular to the axial direction, and “tangential” or “circumferential” refers to a direction mutually perpendicular to the axial and radial directions. As used herein, the terms “forward” or “front” refer to a location relatively upstream in an air flow passing through or around a component, and the terms “aft” or “rear” refer to a location relatively downstream in an air flow passing through or around a component. The direction of this flow is shown by the arrow “F” in
While the illustrated example is a low-bypass turbofan engine with a multistage fan, it will be understood that the principles of the present invention are equally applicable to single-stage fans as well as other types of engines having fans, such as high-bypass turbofans.
As used herein, the term “fan” refers to any apparatus in a turbine engine having a rotor with airfoils operable to produce a fluid flow, where at least a part of the fluid flow discharged from the rotor does not pass through any turbine. Stated another way, at least a part of the fluid flow is only used for thrust and not mechanical energy extraction.
The LPT 22 includes a rotor 54 and variable pitch stators 56. For the purpose of providing additional control of the core engine flow, a variable area nozzle 58 may be provided upstream of the low pressure turbine rotor 22. The cross-sectional flow area to the low pressure turbine rotor 54 may be varied by varying the pitch of the variable area nozzle 58 and the variable pitch stators 56 which vary the back pressure on the high pressure turbine rotor and thereby assist in adjusting the high pressure turbine rotor speed.
Downstream of the core 30, the VABI 24 or other variable mixing device is provided to mix the bypass duct flow with the combustion gases discharged from the LPT 22 in the region designated generally at 60 which also forms the inlet to an augmentor 26.
In the illustrated example, the VABI 24 includes a plurality of rotatable vanes 64 which span a passage 66 in the inner wall 38 separating the bypass duct 40 and the core 30 at a point downstream of the LPT 22. The vanes 64 may be operated by a suitable actuator (not shown). Rotation of the vanes 64 to a near vertical position as seen in
The augmentor 62 is circumscribed by a liner 68 which is spaced apart from the engine outer casing 36 so as to form a passage 70 therebetween. The passage 70 has its inlet disposed approximately coplanar to the inlet of the augmentor 62 such that a portion of the bypass stream is directed into the passage 70 to provide cooling air for the augmentor 62. The outlet of the passage 70 terminates intermediate the augmentor 62 and the variable area converging-diverging exhaust nozzle 28 secured to the aft end of the hour casing 36. The augmentor 62 may be of any type well known in the art. In order to assist in modulating the flow in the bypass duct and core 30, the area of the exhaust nozzle 28 may be varied by suitable variable geometry means such as the illustrated linear actuator 74 controlling a hinged flap assembly 76 to vary the cross-sectional area of the exhaust nozzle 28.
In operation, pressurized air from the HPC 16 is mixed with fuel in the combustor 18 and burned, generating combustion gases. Some work is extracted from these gases by the HPT 20 which drives the compressor 16 via the outer shaft 32. The remainder of the combustion gases are discharged from the core 30 into the LPT 22. The LPT 22 extracts work from the combustion gases and drives the fan 12 through the inner shaft 34. The fan 12 receives inlet airflow from the inlet 42, and thereupon pressurizes the airflow, a portion of which (“core flow”) is delivered to the core 30 and the remainder of which (“bypass flow”) is directed to the bypass duct 40.
The core flow and the bypass flow rejoin in the mixing region 60 Thrust is obtained by the discharge of the mixed flow through the variable area converging-diverging exhaust nozzle 28. When needed, thrust augmentation is provided by introducing fuel into the mixed flow within the augmentor 26 and igniting it.
A normal or nominal operating line represents a locus of operating points on the fan map during normal operation of the engine 10, with no variable-cycle aspects. The operating point of the fan 12 along the operating nominal operating line is determined by fuel flowrate, which is a controllable parameter.
To accommodate various operating requirements, it is possible to change the operating characteristics of the fan 12 and therefore move the operating line from the nominal position on the compressor map.
It will be understood that operation of the VABI 24, the exhaust nozzle 28, and/or other variable-cycle devices have the effect of changing the fan map. For example, during high thrust operation, an open condition of the exhaust nozzle 72 and/or an open position of the VABI 24 drives the fan 12 into a flow exceeding the design condition (“overflow”) reducing fan pressure ratio and a lowering the operating line. This is illustrated in
The VABI 24 and/or nozzle 28 described above are only two examples of “variable-cycle” devices. Any device which is operable to change the back pressure sensed by the fan 12 would have the effect of moving the nominal operating line of the fan map and would therefore be considered a “variable-cycle device”. In the example shown in
It will be understood that some deviation from the nominal operating line is to be expected in some circumstances even without deliberate action. However, as used herein, the term “variable-cycle” implies movement of the operating line from the nominal position deliberately and by a significant amount. For example, using the variable-cycle device, the operating line may be moved (e.g. lowered) from its nominal location by about 5% or more of its nominal distance from the stall line.
Non-limiting examples of variable-cycle devices include: a variable-area bypass injector, a variable-area exhaust, a variable high pressure compressor inter-stage bleed system, a fan having a variable pressure ratio, a fan having the capability of bypassing stages, and a variable area turbine. Multiple engine architectures and configurations can be utilized to achieve variable-cycle capability.
A side effect of lowering the operating line of the fan 12 is to move it towards choke, resulting in a large efficiency penalty. This efficiency penalty can be abated by modification of the fan rotor and/or stator to incorporate “splitters”, examples of which are described below.
The rotor 46 described above defines an annular flowpath surface 82, a small section of which is shown in
The fan blades 44 described above extend from the flowpath surface 82. Each fan blade 44 extends from a root 84 at the flowpath surface 82 to a tip 86 and includes a concave pressure side 88 joined to a convex suction side 90 at a leading edge 92 and a trailing edge 94. As best seen in
The fan blades 44 are uniformly spaced apart around the periphery of the flowpath surface 82. A mean circumferential spacing “s” (see
This reduced solidity can minimize the efficiency loss resulting from “overflowing” the fan 12 in the manner described above. The reduced blade solidity can also have the effect of reducing weight and simplifying manufacturing by minimizing the total number of fan airfoils used in a given rotor stage.
An aerodynamically adverse side effect of reduced blade solidity is to increase the rotor passage flow area between adjacent fan blades 44. This increase in rotor passage through flow area increases the aerodynamic loading level and in turn tends to cause undesirable flow separation on the suction side 90 of the fan blade 44, at the inboard portion near the root 84, also referred to as “hub flow separation”.
To reduce or prevent hub flow separation, the rotor stage 80 may be provided with splitters, or “splittered”. In the illustrated example, an array of splitter blades 144 extend from the flowpath surface 82. One splitter blade 144 is disposed between each pair of fan blades 44. In the circumferential direction, the splitter blades 144 may be located halfway or circumferentially biased between two adjacent fan blades 44. Stated another way, the fan blades 44 and splitter blades 144 alternate around the periphery of the flowpath surface 82. Each splitter blade 144 extends from a root 184 at the flowpath surface 82 to a tip 186, and includes a concave pressure side 188 joined to a convex suction side 190 at a leading edge 192 and a trailing edge 194. As best seen in
The splitter blades 144 function to locally increase the hub solidity of the rotor stage 80 and thereby prevent the above-mentioned flow separation from the fan blades 44. A similar effect could be obtained by simply increasing the number of fan blades 44, and therefore reducing the blade-to-blade spacing. This, however, has the undesirable side effect of increasing aerodynamic surface area frictional losses which would manifest as reduced aerodynamic efficiency and increased rotor weight. Therefore, the dimensions of the splitter blades 144 and their position may be selected to prevent flow separation while minimizing their surface area. The splitter blades 144 are positioned so that their trailing edges 194 are at approximately the same axial position as the trailing edges 184 of the fan blades 44, relative to the rotor 46. this can be seen in
The inner band 245 defines an annular inner flowpath surface 282, and the outer casing 36 defines an annular outer flowpath surface 272. The stator vanes 50 extend between the inner and outer flowpath surfaces 282, 272. Each stator vane 50 extends from a root 284 at the inner flowpath surface 282 to a tip 286 at the outer flowpath surface 272 and includes a concave pressure side 288 joined to a convex suction side 290 at a leading edge 292 and a trailing edge 294. As best seen in
The stator vanes 50 are uniformly spaced apart around the periphery of the inner flowpath surface 282. The stator vanes 50 have a mean circumferential spacing “s”, defined as described above (see
As seen in
In operation, there is a potential for undesirable flow separation on the suction side 290 of the stator vane 50, at the inboard portion near the root 284, also referred to as “hub flow separation”. It also tends to cause undesirable flow separation on the suction side 290 of the stator vane 50, at the outboard portion near the tip 286, also referred to as “case flow separation”. Generally, both of these conditions may be referred to as “endwall separation”.
To counter this adverse side effect, one or both of the inner and outer flowpath surfaces 250, 272 may be provided with an array of splitter vanes. In the example shown in
The splitter vanes 350 function to locally increase the hub solidity of the stator and thereby prevent the above-mentioned flow separation from the stator vanes 50. A similar effect could be obtained by simply increasing the number of stator vanes 50, and therefore reducing the vane-to-vane spacing. This, however, has the undesirable side effect of increasing aerodynamic surface area frictional losses which would manifest as reduced aerodynamic efficiency and increased stator weight. Therefore, the dimensions of the splitter vanes 350 and their position may be selected to prevent flow separation while minimizing their surface area. The splitter vanes 350 are positioned so that their trailing edges 394 are at approximately the same axial position as the trailing edges 294 of the stator vanes 50, relative to the outer flowpath surface 272. This can be seen in
The engine having the fan apparatus described herein with splitter airfoils (splitter blades and/or splitter vanes) has several advantages over the prior art. It increases the endwall solidity level locally, reduces the endwall aerodynamic loading level locally, and suppresses the tendency of the airfoil portion adjacent the endwall to want to separate.
The use of a splittered fan enables variable fan pressure and bypass ratio cycles that will yield reduced engine fuel burn levels. It improves variable-cycle turbine engine performance and enables more efficient operation over wider power ranges and flight regimes. The concept is un-intrusive to implement.
The foregoing has described a gas turbine engine with a splittered fan. All of the features disclosed in this specification (including any accompanying claims, abstract and drawings), and/or all of the steps of any method or process so disclosed, may be combined in any combination, except combinations where at least some of such features and/or steps are mutually exclusive.
Each feature disclosed in this specification (including any accompanying claims, abstract and drawings) may be replaced by alternative features serving the same, equivalent or similar purpose, unless expressly stated otherwise. Thus, unless expressly stated otherwise, each feature disclosed is one example only of a generic series of equivalent or similar features.
The invention is not restricted to the details of the foregoing embodiment(s). The invention extends any novel one, or any novel combination, of the features disclosed in this specification (including any accompanying claims, abstract and drawings), or to any novel one, or any novel combination, of the steps of any method or process so disclosed.
