Patent Description:
Gas turbine engines, such as those which power modern commercial and military aircraft, include a compressor section, combustor section and turbine section arranged longitudinally around the engine centerline so as to provide an annular gas flow path. The compressor section compresses incoming atmospheric gases that are then mixed with a combustible fuel product and burned in the combustor section to produce a high energy exhaust gas stream. The turbine section extracts power from the exhaust gas stream to drive the compressor section. The exhaust gas stream produces forward thrust as it rearwardly exits the turbine section. Some engines may include a fan section, which is also driven by the turbine section, to produce bypass thrust. Downstream of the turbine section, a military engine may include an augmentor section, or "afterburner", that is operable to selectively increase the thrust. The increase in thrust is produced when fuel is injected into the core exhaust gases downstream of the turbine section and burned with the oxygen contained therein to generate a second combustion.

Aircraft engines are sized for a required flight thrust at the most critical conditions, such as top of climb, as well as take-off with one engine failed, or other aircraft flight thrust development margins. Thus, the engine is oversized for a cruise thrust conditions that increases engine/aircraft weight, performance, and fuel consumption penalty accordingly.

Typical Turbofan engines require a fan diameter increase to increase engine Bypass Ratio (BPR) for cruise propulsive efficiency and TSFC improvement, accordingly. Using traditional engine designing approaches, to increase required top of climb thrust for new generation engines with a low fan pressure ratio, the fan diameter would need to be increased by <NUM>-<NUM>% vs. current engines. Such a fan diameter increase, however, requires an increase to engine core size and may increase engine/aircraft installation penalties which then limit TSFC improvement. <CIT> discloses a gas turbine engine of the prior art. A gas turbine engine with a short span afte-fan turbine is known from <CIT>.

An after-fan system for a gas turbine engine according to one aspect of the present invention is provided by claim <NUM>.

An optional embodiment of the present disclosure includes that the variable pitch fan exit guide vane array comprises a split variable pitch fan exit guide vane array.

An optional embodiment of the present disclosure includes that the split variable pitch fan exit guide vane array includes a first variable pitch fan exit guide vane array and a second variable pitch fan exit guide vane array inboard of the first variable pitch fan exit guide vane array.

An optional embodiment of the present disclosure includes that the first variable pitch fan exit guide vane array and the second variable pitch fan exit guide vane array are independently adjustable in pitch.

An optional embodiment of the present disclosure includes that the first variable pitch fan exit guide vane array and the second variable pitch fan exit guide vane array are separated by a splitter.

An optional embodiment of the present disclosure includes that the after-fan turbine is located within the splitter.

An optional embodiment of the present disclosure includes that the after-fan turbine is located downstream of the splitter.

An optional embodiment of the present disclosure includes that the variable pitch fan exit guide vane array is downstream of a fan section.

An optional embodiment of the present disclosure includes that the after-fan turbine is driven in concert with the fan section.

An optional embodiment of the present disclosure includes that the after-fan turbine is driven by a geared architecture which also drives the fan section.

An optional embodiment of the present disclosure includes that the gas turbine engine is a high bypass gas turbine engine.

An optional embodiment of the present disclosure includes that the gas turbine engine is a low bypass gas turbine engine.

An optional embodiment of the present disclosure includes that the low bypass gas turbine engine is variable cycle.

A method of generating thrust for a gas turbine engine according to one aspect of the present invention is provided by claim <NUM>.

An optional embodiment of the present disclosure includes that varying the pitch of the variable fan exit guide vane array comprises independently varying a pitch of an outer diameter section of the variable fan exit guide vane array and an inner diameter section of the variable fan exit guide vane array.

An optional embodiment of the present disclosure includes further comprising locating a splitter between the outer diameter section of the variable fan exit guide vane array and the inner diameter section of the variable fan exit guide vane array.

It should be appreciated; however, the following description and drawings are intended to be exemplary in nature and non-limiting.

The gas turbine engine <NUM> is disclosed herein as a two-spool turbofan that generally incorporates a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM>, and a turbine section <NUM>. The fan section <NUM> drives air along a bypass flowpath B while the compressor section <NUM> drives air along a core flow combustion gas path C for compression and communication into the combustor section <NUM>, then expansion through the turbine section <NUM>. The fan, compressor, and turbine sections may include various architectures that, for example, include a multitude of stages, each with or without various combinations of variable or fixed guide vanes. The sections are defined along a central longitudinal engine axis A. Although depicted as a turbofan in the disclosed non-limiting embodiment, it should be appreciated that the concepts described herein may be applied to other engine architectures such as turbojets, turboshafts, and three-spool (plus fan) turbofans.

The engine <NUM> generally includes a low spool <NUM> and a high spool <NUM> mounted for rotation about an engine central longitudinal axis A relative to an engine static structure <NUM> via several bearings <NUM>. The low spool <NUM> generally includes an inner shaft <NUM> that interconnects a fan section <NUM>, a low pressure compressor ("LPC") <NUM> and a low pressure turbine ("LPT") <NUM>. The inner shaft <NUM> drives the fan section <NUM> directly or through a geared architecture <NUM> that drives the fan section <NUM> at a lower speed than the low spool <NUM>. An exemplary reduction transmission is an epicyclic transmission, such as a planetary or star gear system.

The high spool <NUM> includes an outer shaft <NUM> that interconnects a high pressure compressor ("HPC") <NUM> and high pressure turbine ("HPT") <NUM>. A combustor <NUM> is arranged between the high pressure compressor <NUM> and the high pressure turbine <NUM>. The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes.

Core airflow is compressed by the LPC <NUM>, then the HPC <NUM>, mixed with the fuel and burned in the combustor <NUM>, then expanded over the HPT <NUM> and the LPT <NUM> which rotationally drive the respective high spool <NUM> and the low spool <NUM> in response to the expansion. The main engine shafts <NUM>, <NUM> are supported at a plurality of points by bearings <NUM> within the static structure <NUM>.

In one non-limiting embodiment, the gas turbine engine <NUM> is a high-bypass geared architecture engine in which the bypass ratio is greater than six (<NUM>:<NUM>). The geared architecture <NUM> can include an epicyclic gear train, such as a planetary gear system, star gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than <NUM>, and in another example is greater than <NUM>. The geared turbofan enables operation of the low spool <NUM> at higher speeds which can increase the operational efficiency of the low pressure compressor <NUM> and low pressure turbine <NUM> and render increased pressure in a fewer number of stages.

A pressure ratio associated with the low pressure turbine <NUM> is pressure measured prior to the inlet of the low pressure turbine <NUM> as related to the pressure at the outlet of the low pressure turbine <NUM> prior to an exhaust nozzle of the gas turbine engine <NUM>. In one non-limiting embodiment, the bypass ratio of the gas turbine engine <NUM> is greater than ten (<NUM>:<NUM>), the fan diameter is significantly larger than that of the low pressure compressor <NUM>, and the low pressure turbine <NUM> has a pressure ratio that is greater than five (<NUM>:<NUM>). It should be appreciated, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.

The high bypass ratio results in a significant amount of thrust. The fan section <NUM> of the gas turbine engine <NUM> is designed for a particular flight condition - typically cruise at <NUM> Mach and <NUM>,<NUM> feet. This flight condition, with the gas turbine engine <NUM> at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fan section <NUM> without the use of a fan exit guide vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine <NUM> is less than <NUM>. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of ("T" / <NUM>)<NUM>. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine <NUM> is less than <NUM> fps (<NUM>/s). The Low Corrected Fan Tip Speed in another non-limiting embodiment of the example gas turbine engine <NUM> is less than <NUM> fps (<NUM>/s).

With reference to <FIG>, the fan section <NUM> includes an after-fan system <NUM> downstream of the fan section <NUM>. The after-fan system <NUM> includes a variable pitch fan exit guide vane array <NUM> that form a variable pitch fan exit guide vane ring <NUM> transverse to the bypass flowpath between an outer profile <NUM> of a core static structure <NUM> and an inner periphery <NUM> of the fan nacelle <NUM>. Downstream of the variable pitch fan exit guide vane ring <NUM>, an after-fan turbine <NUM> is positioned between the outer profile <NUM> of the core static structure <NUM> and a ring splitter <NUM> located radially inboard of the inner periphery <NUM> of the fan nacelle <NUM>. In this embodiment, the after-fan turbine <NUM> is a short span after-fan turbine <NUM> which is contained within the ring splitter <NUM>. The ring splitter <NUM> may be at least partially supported by an array of struts <NUM> between the outer profile <NUM> of the core static structure <NUM> and the inner periphery <NUM> of the fan nacelle <NUM>.

The after-fan turbine <NUM> rotates about the engine centerline axis A. The after-fan turbine <NUM> may be driven at speed related to the fan section <NUM> either directly or through the geared architecture <NUM>. The after-fan turbine <NUM> extracts a portion of the energy from the compressed bypass flow from the fan section <NUM> and returns energy to the fan section <NUM> through torque. The after-fan turbine <NUM> includes after-fan turbine blades <NUM> that, in this embodiment, are of a span less than the fan blades <NUM> of the fan section <NUM>. In one example, the after-fan turbine <NUM> is of a diameter between <NUM>%-<NUM>% of the fan section <NUM>. The stagger angles of the after-fan turbine blades <NUM> and fan blades <NUM> are the angle of the chord line connecting the tip of the airfoil shape to the trailing edge of the airfoil shape. Measured between the direction of the wheel speed, U, and the chord line, the stagger angles of the after-fan turbine blades <NUM> and fan blades <NUM> are substantially the same. The chord length of the fan blades <NUM> is substantially longer than the chord length of the after-fan turbine blades <NUM> due to the different aerodynamic loadings of the fan blades <NUM> versus the after-fan turbine blades <NUM>. The relative velocity, W, of the flow passing over the fan blades <NUM> decelerates and creates a higher aerodynamic loading. The relative velocity, W, of the flow passing over after-fan turbine blades <NUM> accelerates from W1 to W2 and the aerodynamic loading is lower than the fan blades <NUM>. The aerodynamic loading of the after-fan turbine blades <NUM> varies with the vane <NUM> angle, alpha1. Decreasing alpha1 increases the aerodynamic loading of the after-fan turbine blades <NUM>.

A pitch angle of each of the variable pitch fan exit guide vane array <NUM> may be varied along a pitch axis V to change the pitch thereof in response to a controller <NUM> to modify the bypass airflow from the fan section <NUM> that is communicated into the after-fan turbine <NUM>. The controller <NUM> generally may include a processor, a memory, and an interface. The processor may be any type of microprocessor having desired performance characteristics. The processor and the interface are communicatively coupled to the memory. The memory may be embodied as any type of computer memory device which stores data and control algorithms such as logic as described herein. The interface is communicatively coupled to a number of hardware, firmware, and/or software components, including sensors and actuators <NUM> for the variable pitch fan exit guide vane array <NUM>. The controller <NUM> may, for example, be a portion of a flight control computer, a portion of a Full Authority Digital Engine Control (FADEC), a stand-alone unit or combinations thereof.

The variable pitch fan exit guide vane array <NUM> modifies the bypass airflow from the fan section <NUM> (<FIG>) such that a bypass airflow entrance angle into the after-fan turbine <NUM> is selectively adjusted to change the energy extracted therefrom to, for example, maximize energy extraction at cruise, yet minimize energy extraction at the top of climb and take-off thrust conditions.

The variable pitch fan exit guide vane array <NUM> and the after-fan turbine <NUM> controls a fan operating line for efficiency and operability at the desired flight conditions. The after-fan turbine <NUM> optimizes fan duct nozzle <NUM> pressure ratio, and nozzle <NUM> exit airflow velocity, accordingly, to increase propulsive efficiency at different off-design flight conditions. The short span after-fan system <NUM> is of reduced weight and provides design flexibility to create an aerodynamically efficient after fan turbine <NUM> by design choices of rpm and diameter optimization for achieving engine thrust requirements.

To decrease engine specific fuel consumption (TSFC) engine propulsive efficiency is increased, which is associated with very low cruise fan pressure ratio of about <NUM>, and a super high engine bypass ratio of about <NUM> vs. current industry cruise at a fan pressure ratio of about <NUM> and engine bypass ratio of about <NUM>. Current engines have a top of climb (for max available thrust) fan pressure ratio of about <NUM>.

In one embodiment, the variable pitch fan exit guide vane array <NUM> are movable (<FIG>) between a maximum power extraction pitch position <NUM> for the after-fan system <NUM> typically utilized at cruise thrust conditions and a minimum power extraction pitch position <NUM> typically utilized for max climb engine power setting, top of climb, and at take-off thrust conditions as represented by the airflow vectors triangles. The after-fan system <NUM> permits the fan diameter of the fan section <NUM> to be sized for cruise thrust conditions which results in one example of an about <NUM>% TSFC advantage. The after-fan system <NUM> also permits an increase to the max available fan pressure ratio for the top of climb, and at take-off thrust conditions. This provides a decrease in overall engine weight with a reduced diameter to facilitate underwing nacelle installation. In this embodiment, the fan section <NUM> and the after-fan turbine <NUM> rotate in the same direction, however, counter-rotating systems may also be provided.

With reference to <FIG>, example ranges of the airflow vector angles in the airflow vectors triangles include:.

These ranges may vary depending on, for example, turbine blade section arrangements along turbine vane span, after-fan blade geometry, anticipated turbine expansion ratios for the embodiment, turbine rpm, and turbine blade profiles. For this high swirl version fan, the existed high fan airflow swirl create condition to use the swirl to improve variable vane efficiency and design when using contra rotating version. For contra rotating versions, vane and turbine blades profiles will have mirrored arrangement vs. considered in <FIG> vector triangles. Further, the after-fan turbine blades may have different aerodynamic loading profiles along blades span with a minimum turbine expansion ratio in a root portion of the blade, which faces core flow path.

With reference to <FIG>, another embodiment of the after-fan system 60A includes a variable pitch fan exit guide vane ring 64A upstream of the after-fan turbine 74A. The variable pitch fan exit guide vane ring 64A includes an outer diameter portion 62A of the variable pitch fan exit guide vane ring 64A and an inner diameter portion 62B of the variable pitch fan exit guide vane ring 64A. The outer diameter portion 62A and the inner diameter portion 62B may be independently adjusted in pitch about the common pitch axis V by a respective actuator 84A, 84B. Alternatively, the outer diameter portion 62A or the inner diameter portion 62B may be fixed to at least partially supports the ring splitter 76A. Alternatively, the ring splitter 76A may be partially supported by a strut array <NUM>. In this embodiment, the after-fan turbine <NUM> is contained within the ring splitter 76A. This enables the portion of the bypass flow B in the outer diameter portion 62A to be adjusted in flow rate and pressure independently of the portion of the bypass flow B in the inner diameter portion 62B in response to engine <NUM> fuel flow changes that vary engine <NUM> thrust as between a cruise and climb condition. The short blade span turbine splitter provides advantage for turbine blade tip clearance to facilitate efficiency.

With reference to <FIG>, another embodiment of the after-fan system 60D is utilized with a low-bypass gas turbine engine <NUM>. The gas turbine engine <NUM> is disclosed herein as a two-spool, low-bypass, augmented turbofan that generally incorporates a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM>, a turbine section <NUM>, an augmenter section <NUM>, a duct section <NUM>, and a nozzle system <NUM> along a central longitudinal engine axis A. Although generally described and illustrated with regard to a low-bypass gas turbine engine, a variable cycle gas turbine engine that essentially alters a bypass ratio during flight to achieve countervailing objectives such as high specific thrust for high-energy maneuvers yet optimizes fuel efficiency for cruise and loiter operational modes will also benefit herefrom.

An outer case structure <NUM> and an inner case structure <NUM> define a generally annular secondary airflow path <NUM> around a core airflow path <NUM>. Various structures may define the outer case structure <NUM> and the inner case structure <NUM> which essentially define an exoskeleton to support rotational hardware therein. Air that enters the fan section <NUM> is divided between core airflow through the core airflow path <NUM>, and secondary airflow through the secondary airflow path <NUM>. The core airflow passes through the combustor section <NUM>, the turbine section <NUM>, then the augmentor section <NUM>, where fuel may be selectively injected and burned to generate additional thrust through the nozzle system <NUM>.

The secondary airflow may be utilized for a multiple of purposes to include, for example, cooling, pressurization and variable cycle operations. The secondary airflow as defined herein is any airflow different from the core airflow. The secondary airflow may ultimately be at least partially injected into the core airflow path <NUM> adjacent to the duct section <NUM> and the nozzle system <NUM>.

With reference to <FIG>, in this embodiment, the after-fan system 60D includes a split variable pitch fan exit guide vane ring 164D upstream of an after-fan turbine 174D that is upstream of the secondary airflow path <NUM> and the core airflow path <NUM>. The split variable pitch fan exit guide vane ring 164D includes an outer diameter portion of the variable pitch fan exit guide 162A and a second vane portion 162B that partially supports the ring splitter 176D. The outer diameter portion of the variable pitch fan exit guide 162A and/or the second vane portion 162B may be individually or collectively varied in pitch to selectively control the airflow to the after-fan turbine 174D within the ring splitter 176D. The ring splitter 176A is located radially intermediate the secondary airflow path <NUM>. Architectures with two portion variable pitch exit guide vanes that facilitate control of turbine power extraction by inner portion airflow vector changes in front of the turbine blade leading edge, and at the same time controls a fan operating line.

With reference to <FIG>, another embodiment of the after-fan system 60F includes a split variable pitch fan exit guide vane ring 164F upstream of a short span after-fan turbine 174F that is upstream of the secondary airflow path <NUM> and the core airflow path <NUM>. The split variable pitch fan exit guide vane ring 164F includes an outer diameter portion 162Fa and a inner diameter portion 162Fb that partially supports the ring splitter 176F. The outer diameter portion 162Fa and/or the inner diameter portion 162Fb may be individually or collectively varied in pitch to selectively control the airflow to the after-fan turbine 174F within the ring splitter 176F. In this embodiment, the ring splitter 176F is located radially intermediate the secondary airflow path <NUM> and encompasses a low pressure compressor 124A of the compressor section <NUM>. That is, the fan section <NUM> and the split variable pitch fan exit guide vane ring 164F extend radially beyond the ring splitter 176F. The ring splitter 176F is also at least partially supported by a strut array <NUM> which spans the secondary airflow path <NUM> and the core airflow path <NUM> to at least partially support the inner case structure <NUM>.

The split variable pitch fan exit guide vane ring 164F permits independent control of the outer diameter portion 162Fa and the second vane portion 162Fb. The outer diameter portion 162Fa facilitates control of the fan operational line for best efficiency and stability margin, and at the inner diameter portion 162Fb facilitates the required power extraction and efficiency for the short span after-fan turbine 174F. The outer diameter portion 162Fa also facilitates a change in an adaptive fan outer airflow to optimized BPR for different flight segments. The axially extended ring splitter 176F facilitates formation of a third stream by splitter extension downstream of the fan duct. Architectures with short span blades facilitates the generation of the third stream by splitter extension downstream of the duct to optimize engine bypass ratio for different flights segments.

Claim 1:
An after-fan system for a gas turbine engine, comprising:
a variable pitch fan exit guide vane array (<NUM>); and a control (<NUM>) operable to vary a pitch of the variable fan exit guide vane array (<NUM>);
characterized in comprising
a short span after-fan turbine (<NUM>) downstream of the variable pitch fan exit guide vane array (<NUM>) and upstream of a bypass or secondary airflow path (B, <NUM>) and a core airflow path (C, <NUM>), the short span after-fan turbine (<NUM>) being positioned between an outer profile (<NUM>) of a core static structure (<NUM>) and a ring splitter (<NUM>) located radially inboard of an inner periphery (<NUM>) of a fan nacelle or outer case structure (<NUM>, <NUM>), the short span after-fan turbine (<NUM>) being contained radially within the ring splitter (<NUM>).