Gas turbine engine with a multi-spool driven fan

A gas turbine engine includes low and high spools constructed and arranged to rotate about an engine axis. The low spool drives at least one leading stage of a fan section and the high spool drives an aft stage of the fan section. The aft stage may generally include a bypass duct for controllably flowing a bypass stream directly from the leading stage and controllably and/or selectively into an auxiliary flowpath and/or into a second flowpath both located radially outward from a core flowpath.

BACKGROUND

The present disclosure relates to a gas turbine engine, and more particularly to an engine having a multi-spool driven fan section.

Gas turbine engines, such as those which power modern military aircraft, include a compressor section to pressurize a supply of air, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases and generate thrust. Downstream of the turbine section, an augmentor section, or “afterburner”, 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.

Variable cycle gas turbine engines power aircraft over a range of operating conditions yet achieve countervailing objectives such as high specific thrust and low fuel consumption. The variable cycle gas turbine engine essentially alters a bypass ratio during flight to match requirements. This facilitates efficient performance over a broad range of altitudes and flight conditions to generate high thrust for high-energy maneuvers yet optimize fuel efficiency for cruise and loiter.

Variable cycle gas turbine engines require an effective actuation system to vary the bypass ratio (BPR) between maximum afterburning conditions and cruise conditions to operate the engine at various cycle points. Variable cycle gas turbine engines are typically of a three-stream engine architecture in which a two-stage fan directly feeds all three streams, e.g., core stream, second stream, third stream. Typically, a part-span booster fan stage feeds the core stream and the second stream. Although effective, this architecture requires a relatively complex fan design and a challenging intermediate case design due to the limited area available to execute three streams in the same required package of traditional two stream engines.

SUMMARY

A gas turbine engine according to one, non-limiting, embodiment of the present disclosure includes a high spool constructed and arranged to rotate about an engine axis; a low spool constructed and arranged to rotate about the engine axis; a fan leading stage connected for rotation to the low spool; and a fan aft stage connected for rotation to the high spool.

Additionally to the foregoing embodiment, the engine includes a high pressure turbine connected to the high spool.

In the alternative or additionally thereto, in the foregoing embodiment, the engine includes a low pressure turbine connected to the low spool.

In the alternative or additionally thereto, in the foregoing embodiment, the engine includes a fan outer housing axially aligned to and circumferentially surrounding the leading and aft stages; and a fan inner housing spaced radially inward from the outer housing, located aft of the leading stage, and circumferentially surrounding the aft stage, wherein a bypass duct is defined radially between the outer and inner housing for at least partially flowing air from the leading stage.

In the alternative or additionally thereto, in the foregoing embodiment, the engine includes a core case concentrically located about the engine axis, wherein a core flowpath is defined radially inward of the core case for flowing a core stream; a second case spaced radially outward of the core case, wherein a second flowpath is defined radially between the core and second case for flowing a second stream; and an outer case spaced at least in-part radially outward from the second case, wherein an auxiliary flowpath is defined at least in-part radially between the second and outer cases for flowing an auxiliary stream, and the bypass duct is in selective, direct, communication with at least one of the second and auxiliary flowpaths.

In the alternative or additionally thereto, in the foregoing embodiment, the engine includes a control valve arrangement constructed and arranged to control airflow from the bypass duct to the second flowpath.

In the alternative or additionally thereto, in the foregoing embodiment, the engine includes a control valve arrangement constructed and arranged to control airflow from the bypass duct to the auxiliary flowpath.

In the alternative or additionally thereto, in the foregoing embodiment, the control valve arrangement includes an isolation valve feature to selectively prevent flow through the bypass duct.

In the alternative or additionally thereto, in the foregoing embodiment the control valve arrangement includes a selector valve feature for controllably diverting airflow between the auxiliary and second flowpaths.

In the alternative or additionally thereto, in the foregoing embodiment, the engine includes an airflow control system including at least one heat exchanging device arranged to heat a portion of the second stream from the second flowpath upon entry into the auxiliary flowpath.

In the alternative or additionally thereto, in the foregoing embodiment, the at least one heat exchanging device includes a heat exchanger and a control valve constructed and arranged to control flow through the heat exchanger.

In the alternative or additionally thereto, in the foregoing embodiment, the engine includes a control valve arrangement constructed and arranged to control airflow from the bypass duct to the second flowpath, and wherein the airflow control system includes a controller constructed and arranged to control the control valve arrangement and the control valve of the heat exchanging device.

In the alternative or additionally thereto, in the foregoing embodiment, the heat exchanging device includes a conduit loop in direct fluid communication between the core flowpath and a hot side of the heat exchanger for flowing heated air from the core flowpath to the hot side and returning the cooled hot air from the hot side and through the core case for cooling components within the core case, and wherein a cold side of the heat exchanger is constructed and arranged to flow the portion of the second stream for cooling the hot air.

A gas turbine engine according to another, non-limiting, embodiment includes a low pressure turbine; a high pressure turbine proximate to the low pressure turbine; a first fan stage driven by the low pressure turbine; and a second fan stage driven by the high pressure turbine.

Additionally to the foregoing embodiment, the second fan stage includes a bypass duct in direct, selective, communication between an auxiliary flowpath and a second flowpath.

A method of operating a gas turbine engine according to another, non-limiting, embodiment, includes the steps of driving a leading stage of a fan section with a low spool; and driving an aft stage of a fan section with a high spool.

Additionally to the foregoing embodiment, the method includes the steps of flowing an incoming airflow through the leading stage; and utilizing a flow control system to controllably flow at least a portion of the incoming airflow through the aft stage.

In the alternative or additionally thereto, in the foregoing embodiment, the method includes the step of flowing a portion of the incoming airflow through a bypass duct that bypasses the aft stage.

In the alternative or additionally thereto, in the foregoing embodiment, the method includes the step of controlling a valve arrangement with a controller to control the flow of a bypass stream through the bypass duct.

In the alternative or additionally thereto, in the foregoing embodiment, the method includes the step of controlling the valve arrangement to selectively expel the bypass stream into a second stream, an auxiliary stream, or both.

DETAILED DESCRIPTION

FIG. 1schematically illustrates a gas turbine engine20. The gas turbine engine20is disclosed herein as a two-spool turbo fan that generally incorporates a fan section22, a compressor section24, a combustor section26and a turbine section28. Alternative engines might include a turbine exhaust case section30followed in a downstream direction by an augmentor section32, an exhaust duct section34and a nozzle section36. The fan section22drives air along a second flowpath while the compressor section24drives air along a core flowpath for compression and communication into the combustor section26then expansion through the turbine section28. The fan, compressor and turbine sections22,24,28may each include various architectures that, for example, include a plurality of stages, each with or without various combinations of blades and variable or fixed guide vanes orientated about an engine axis A.

Variable cycle gas turbine engines power aircraft over a range of operating conditions and essentially alters a bypass ratio during flight to achieve objectives that may be countervailing (such as high specific thrust for high-energy maneuvers) yet optimizes fuel efficiency for cruise and loiter modes of operation. Although depicted as an augmented low bypass turbofan in the disclosed, non-limiting, embodiment, it is understood that the concepts described herein are applicable to other types of turbine engines including non-augmented engines, geared architecture engines, high bypass and/or direct drive turbofans, turboshafts, and three-spool (plus fan) turbofans wherein an intermediate spool includes an intermediate pressure compressor (“IPC”) between a Low Pressure Compressor (“LPC”) and a High Pressure Compressor (“HPC”), and an Intermediate Pressure Turbine (“IPT”) between the High Pressure Turbine (“HPT”) and the Low Pressure Turbine (“LPT”).

The engine20generally includes a low spool38and a high spool40mounted for rotation about the engine axis A, and relative to an engine core case42, via several bearing structures (not shown). The low spool38generally includes an inner shaft that interconnects at least one leading fan stage44of the fan section22, a LPC46of the compressor section24and a LPT48of the turbine section28. The inner shaft of the low spool38drives the leading fan stage44directly, or through a geared architecture50to drive the leading fan stage44at a lower speed than the low spool38. An exemplary reduction transmission is an epicyclic transmission of the architecture50, namely a planetary or star gear system.

The high spool40includes an outer shaft that interconnects at least one aft fan stage52of the fan section22, a HPC54of the compressor section24and a HPT56of the turbine section28. The outer shaft of the high spool40drives the aft fan stage52directly, or through a geared architecture58to drive the aft fan stage52at a lower speed than the high spool40. Like the geared architecture50, an exemplary reduction transmission of the architecture58is an epicyclic transmission, namely a planetary or star gear system.

A combustor (not shown) of the combustor section26is arranged between the HPC54and the HPT56and, at least in-part, radially within a diffuse case module60of the core engine case42. The inner and outer shafts of the respective low and high spools38,40may be concentric and rotate about the engine axis A that is collinear with their, respective, longitudinal axis. A core air stream (see arrow62) is compressed by the LPC46then the HPC54, mixed with fuel and burned in the combustor of the combustor section26, then expanded over the HPT56and the LPT48. The LPT48and HPT56rotationally drive the respective low spool38and high spool40in response to the expansion.

By tying the at least one aft stage52of the fan section22to the HPT56, the fan pressure ratio will lapse more strongly from high engine power to low engine power (i.e. steeper operating characteristic); thereby, improving part power propulsion efficiency and reducing fuel consumption at cruise. Also, by distributing the fan section22work across two spools38,40, the turbomachinery will be more aerodynamically and thermodynamically balanced allowing for a more optimal HPT and LPT component aero design and lighter weight.

In one non-limiting example, the gas turbine engine20is a high-bypass geared aircraft engine. In a further example, the gas turbine engine20bypass ratio is greater than about six (6:1). The geared architecture48can include an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about 2.3:1, and in another example is greater than about 2.5:1. The geared turbofan enables operation of the low spool38at higher speeds that can increase the operational efficiency of the LPC46and LPT48and render increased pressure in a fewer number of stages.

A pressure ratio associated with the LPT48is pressure measured prior to the inlet of the LPT48as related to the pressure at the outlet of the LPT48prior to the exhaust section36of the gas turbine engine20. In one non-limiting embodiment, the bypass ratio of the gas turbine engine20is greater than about ten (10:1), the fan diameter is significantly larger than that of the LPC46, and the LPT48has a pressure ratio that is greater than about five (5:1). It should be understood; 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.

In one embodiment, a significant amount of thrust is provided by a bypass flow path due to the high bypass ratio. The fan section22of the gas turbine engine20is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). This flight condition, with the gas turbine engine20at its best fuel consumption, is also known as 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 section22without 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 engine20is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of (T/518.7)0.5in which “T” represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine20is less than about 1,150 feet per second (351 meters per second).

Aside from the core case42, the engine20may include: a fan duct, intermediate, case64spaced radially outward and, at least in-part, axially aligned to the core case42; and an outer case66spaced, at least in-part, radially outward from, and axially aligned to, the intermediate case64. The core case42generally defines a radially outward boundary of an annular core flowpath68for the flow of the core stream62. The core and intermediate cases42,64generally define, respective, radially inward and outward boundaries of an annular second flowpath70for the flow of a second stream (see arrow72) of air. Generally, the second flowpath70may be part of a fan duct and the second stream72may be a high or low bypass air stream depending upon the engine application. The intermediate and outer cases64,66generally define, respective, radially inward and outward boundaries of an annular auxiliary flowpath74for the flow of an auxiliary stream (see arrow76) of air. It is further understood and contemplated that the auxiliary flowpath74may not be annular, and instead may generally be a series of axially extending flowpaths extending axially and circumferentially spaced or segmented from one-another.

A primary stream (see arrow73) of air that enters the aft stage52of the fan section22is divided between the core stream62through the core flowpath68, and the second stream72through the second flowpath70. The core stream62is compressed by the compressor section24mixed with fuel and burned in the combustor section26, then expanded through the turbine section28, at least partially deswirled by airfoils (not shown) of the turbine exhaust case section30, then exhausted through the exhaust duct section34and a mixed flow exhaust nozzle77of the nozzle section36. Fuel may also be selectively injected into the core stream70in the augmentor section32downstream of the turbine section28to generate additional thrust through the mixed flow exhaust nozzle77from the exhaust duct section34.

The exhaust duct section34may be circular in cross-section as typical of an axis-symmetric augmented low bypass turbofan. Alternatively or additionally, the exhaust duct section34may be non-axisymmetric in cross-section or other shape and/or non-linear with respect to the central longitudinal engine axis A to form, for example, a serpentine shape to block direct view to the turbine section. The core flowpath68terminates with the mixed flow exhaust nozzle77of the nozzle section36that may include, for example, various fixed, variable, convergent/divergent, two-dimensional and three-dimensional nozzle systems. The auxiliary flowpath74terminates at an auxiliary nozzle79of the nozzle section36, and generally located radially outward from the exhaust nozzle77.

The second stream72flowing through the second flowpath70may be utilized, for example, to enhance heat transfer or pressurize, another component or cavity. Moreover, the second stream72may be, at least partially, injected into the core flowpath68adjacent the augmentor section32and the exhaust duct section34for exhaust through the mixed flow exhaust nozzle77(illustrated schematically by flow arrows81through an augmentor liner83). That is, the core stream62and the second stream72may be combined and mixed for ejection through the mixed flow exhaust nozzle77.

Referring toFIGS. 1 and 2, the fan section22of the engine20includes inner and outer fan housings78,80. The inner fan housing78may generally be a forward extension of the intermediate case64and is located radially outward from and, at least in-part, axially aligned to the aft fan stage52of the fan section22. The inner housing78may include a forward, distal, edge78E that is circumferentially continuous and located immediately aft of the leading fan stage(s)44. The fan outer housing80is circumferentially continuous, surrounds the leading and aft stages44,52of the fan section22and may generally be a forward projecting extension of the outer case66. A rearward or downstream portion of the outer housing80is spaced radially outward from the inner housing78and, together, form or define an annular bypass duct or flowpath82for the flow of a bypass stream (see arrow84) of air that generally flows from the leading stage44, bypasses the aft stage52and controllably enters the second and/or auxiliary flowpaths70,74to combine with respective air streams72,76.

The aft stage52is located in a primary duct or flowpath85generally defined by a radially inward facing side of the inner housing78. At least a substantial portion of the air that exits the fan leading stage44flows through the primary duct85(and consequently the aft stage52) as the primary airstream73. Immediately downstream of the aft fan stage52, the primary stream73is generally divided into the second stream72and the core stream62.

An airflow control system86of the engine20may facilitate the control of air flow through the bypass duct82, through the auxiliary flowpath74, and through the second flowpath70, or any combination thereof. The control system86may include a controller88and a control valve arrangement90(seeFIG. 2) that may have an isolation valve or duct blocker92, a diverter or selector valve94. The controller88may receive sensory or operator input, processes the input, and outputs electronic control signals96,98that operate the respective isolation and selector valves92,94. It is contemplated and understood that the control valve arrangement90may contain any variety of valve configurations including single-bodied units capable of both pathway isolation and flow diversion between two downstream pathways. It is further understood that the term ‘valve’ may include any variety of structures including a plurality of blocker doors and/or aerodynamically shaped flaps that may be circumferentially distributed and operated by a common sync-ring as one, non-limiting, example.

The controller88may be, for example, part of a flight control computer, a portion of a Full Authority Digital Engine Control (FADEC), a stand-alone unit or other system. The control module typically includes a processor, a memory, and an interface. The processor may be any type of microprocessor having desired performance characteristics. The memory may be any computer readable medium that stores data and control algorithms such as logic as described herein. The interface facilitates communication with other components such as an actuator of the valve arrangement90that operates one or both of the valves92,94.

The isolation valve92may generally be located in the bypass duct82and may be generally supported by at least one of the inner and outer housings78,80. The diverter valve94may generally be integrated between, and supported by at least one of, the inner housing78and the intermediate case64. As best shown inFIG. 2, when the isolation valve92is in a closed position100and the diverter valve94is in a neutral state102that facilitates aerodynamic considerations, all of the incoming airflow (see arrow104inFIG. 1) is directed through both the leading and aft stages44,52of the fan section22. When exiting the fan section22, the airflow104(i.e. with valve92closed) is generally divided into the core stream62and the second stream72. In this configuration, the engine20is capable of maximum power with high thrust that may be advantageous for takeoff flight conditions.

Referring toFIG. 3, when the isolation valve92is in the closed position100and the diverter valve94is in a second stream receiving state106, all of the incoming airflow104is directed through both the leading and aft stages44,52of the fan section22. In addition, a portion (see arrow108) of the second stream72is diverted from the second flowpath70and enters the auxiliary flowpath74to join with the auxiliary stream76. In this configuration, the engine20is capable of operating in an exhaust management mode with high thrust that may be advantageous for takeoff, climb, and high-speed cruise flight conditions.

Referring toFIG. 4, when the isolation valve92is in an open position110and the diverter valve94is in the neutral state102, a portion (i.e. bypass airstream84) of the incoming airflow104that flows through the leading fan stage(s)44does not flow through the aft stage52and instead is bypassed and directly enters the auxiliary flowpath74. Because the diverter valve94is in the neutral position102, none of the second stream72enters the auxiliary flowpath74via the diverter valve94of the valve arrangement90. In this configuration, the engine20is capable of high fuel efficiency operation during loiter and cruise flight conditions with low thrust. That is, the augmented engine20may more closely resemble a high bypass turbofan engine more common in the commercial aviation industry.

Referring toFIG. 5, when the isolation valve92is in the open position110and the diverter valve94is in a controllably adjustable bypass stream receiving state112, at least a portion (see arrow114) of the bypass stream84enters the second flowpath70to join with the second stream72. In addition, the diverter valve94may be appropriately positioned allowing for the remaining portion (see arrow116) of the bypass stream84to enter the auxiliary flowpath74and join with the auxiliary stream76. Although total pressure may be greater in the second pathway70than in the bypass duct82, the static pressure may be about the same, thus allowing controlled flow to occur from the bypass duct82to the second flowpath70. In this configuration, the engine20is capable of operating in an exhaust management mode at low-speed cruise flight conditions with low thrust.

Referring toFIGS. 1 and 6, the airflow control system86may also include at least one heat exchanging device118each having a heat exchanger120, a control valve122, and a pathway or conduit loop124for the supply flow of heated air and subsequent return of cooled air. Each valve122may be controlled via an electronic signal126from the controller88. The heat exchanger120and valve122may be of an annular architecture and/or multiple discrete passages, ducts, or other selectively controlled flow path configurations distributed about the circumference of the second flowpath70. The valve122selectively flows a portion (see arrow128) of the second stream72from the second flowpath70into the auxiliary flowpath74via the cold side of the heat exchanger120for cooling the air flowing through pathway124. With the control valve(s)122open, the stream portion128is generally heated and becomes part of the auxiliary stream76thereby acting as a ‘heat sink’ and increasing the auxiliary stream temperature.

As best shown inFIG. 6, examples of heat exchanging devices may include a device118A with a pathway124A that receives hot air from within the diffuser case module60and returns the air in a cooled state to the compressor section24for cooling of components therein. As another example, a device118B may have a pathway124B that receives hot air from the turbine section28and returns the air in a cooled state to the turbine section28for cooling of components therein. Yet another example may include a device118C that may have a pathway124C that receives hot air from a component130that is external to the engine20and returns the air to a location external from the engine20. It is further contemplated and understood that the other pathways may flow hot air sourced from other engine sections, stages, or systems, and return the air in a cooled state to any one or combination of the engine sections, stages, or systems.

The heat exchanging device(s)118may also be utilized to vary the pressure drop in the second flowpath70, as the pressure drop through the cold side of the heat exchanger120is enhanced through discharge of the selected portion128of the second stream72into the auxiliary flowpath74, which may be at a lower pressure. It should also be appreciated that although particular systems are separately defined and schematically illustrated, such as the heat exchanger120and the valve122, each, or any, may be otherwise combined or segregated. Alternatively, an operable valve may not be used and instead, the heat exchanger itself may have a pre-determined pressure drop (e.g. pre-sized, internal orifices) capable of establishing flows within an acceptable range. Yet further, the stream portion128flow may, at least in-part, be controlled via actuation of the valve arrangement90(i.e. with or without use of valve122) that may controllably vary the pressure differential between the second flowpath70and the auxiliary flowpath74.

Referring toFIG. 7, the engine20may operate under varying modes132and flight conditions134with a desired engine thrust level136. For instance, with a mode132of maximum power and a flight condition134of takeoff, the thrust level136may be high. With this configuration, the airflow control system86may orientate the isolation valve92in the closed position100and the selector valve94in the neutral state102. With this orientation, the second stream portion128flowing through the heat exchanger120of the heat exchanging device(s)118is not sourced from the bypass duct82and is solely sourced from the primary duct85. The auxiliary stream76flow is thus sourced solely from the heat exchanger120cold side (i.e. second stream portion128flow) with no flow attributed directly from the bypass duct82or the primary duct85. This results in a low flow condition through the auxiliary nozzle79. The second stream72flow sources are attributable directly from the primary duct85with no flow directly from the bypass duct82, resulting in a maximum flow condition through the second stream nozzle77.

When the engine20operates with the mode132of exhaust management and a flight condition134of takeoff, climb and/or high cruise, the thrust level136may be high. With this configuration, the airflow control system86may orientate the isolation valve92in the closed position100and the selector valve94in the selected state106. With this orientation, the second stream portion128flowing through the heat exchanger120of the heat exchanging device(s)118is not sourced from the bypass duct82and is solely sourced from the primary duct85. The auxiliary stream76flow is thus sourced mostly from the primary duct85with a minor portion from the heat exchanger120cold side (i.e. second stream portion128flow), and with no flow attributed directly from the bypass duct82. This results in a medium flow condition through the auxiliary nozzle79. The second stream72flow sources are attributable directly from the primary duct85with no flow directly from the bypass duct82, resulting in a high flow condition through the second stream nozzle77.

When the engine20operates with the mode132of high efficiency and a flight condition134of loiter or cruise, the thrust level136may be medium. With this configuration, the airflow control system86may orientate the isolation valve92in the open position110and the selector valve94in the neutral state102. With this orientation, the second stream portion128flowing through the heat exchanger120of the heat exchanging device(s)118is not sourced from the bypass duct82and is solely sourced from the primary duct85. The auxiliary stream76flow is thus sourced mostly and directly from the bypass duct82with a minor portion from the heat exchanger120cold side (i.e. second stream portion128flow), and with no flow attributed directly from the primary duct85. This results in a high flow condition through the auxiliary nozzle79. The second stream72flow sources are attributable directly from the primary duct85with no flow directly from the bypass duct82, resulting in a medium flow condition through the second stream nozzle77.

When the engine20operates with the mode132of exhaust management and a flight condition134of low-speed cruise, the thrust level136may be low. With this configuration, the airflow control system86may orientate the isolation valve92in the open position110and the selector valve94in the selected state112. With this orientation, the second stream portion128flowing through the heat exchanger120of the heat exchanging device(s)118is sourced mostly and directly from the primary duct85with a minor portion sourced directly from the bypass duct82. The auxiliary stream76flow is thus sourced mostly and directly from the bypass duct82with a minor portion from the heat exchanger120cold side (i.e., second stream portion128flow), and with no flow attributed directly from the primary duct85. This results in a medium flow condition through the auxiliary nozzle79. The second stream72flow sources are mostly attributable directly from the primary duct85with a minor portion attributable directly from the bypass duct82, resulting in a high flow condition through the second stream nozzle77.

With the high spool driven aft stage52and the valve arrangement90, the fan section22is higher in efficiency and higher in maximum fan pressure ratio (FPR) capability without compromise, when compare to more traditional engines. In addition, the HPT56is higher in efficiency and requires less cooling, and the LPT48may employ fewer stages and requires less cooling than more traditional engines. Similar to more traditional three stream engines, the present engine20with the HPT driven fan stage52provides for cruise power fan flow holding for reduced spillage drag. The auxiliary stream76pressure may be similar to a stage three in the more traditional three stream engines because of the pressure drop across the heat exchanging devices118resulting in similar propulsion efficiencies. The engine20of the present disclosure has higher fan pressure ratio capability than more traditional engines with a more balanced turbomachinery. This provides for a more compact propulsion system that is lighter and shorter.

While the invention is 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 without departing from the spirit and scope of the invention. In addition, different modifications may be made to adapt the teachings of the invention to particular situations or materials, without departing from the essential scope thereof. The invention is thus not limited to the particular examples disclosed herein, but includes all embodiments falling within the scope of the appended claims.