JP6059861B2 - Gas turbine engine and method for operating a gas turbine engine - Google Patents

Description

  The present invention relates generally to gas turbine engines, and more specifically to a gas turbine engine having an intermediate pressure fan stage having a tip fan installed on a blade driven by a pressurized air flow.

  In a turbofan aircraft gas turbine engine, during operation, air is pressurized in the fan module, optional booster module, and compressor module. A portion of the air flowing through the fan module flows as a bypass stream and is used to generate a portion of the thrust necessary to propel the aircraft in flight. Air sent through the optional booster module and compressor module is mixed with fuel in the combustor and ignited to generate hot combustion gas, which flows through the turbine stage, The energy is extracted from the hot combustion gases to power the fan, booster and compressor rotor. The fan, booster and compressor module has a series of rotor stages and stator stages. The fan and booster rotor are typically driven by a low pressure turbine (LPT), and the compressor rotor is driven by a high pressure turbine (HPT). Although the fan and booster rotor are aerodynamically coupled to the compressor rotor, the fan rotor and compressor rotor typically operate at different machine rotational speeds.

  In many cases, components of lower bypass ratio engines with higher fan pressure ratios include compressors, combustors, high pressure turbines (HPT) and high bypass commercial or intermediate bypass engines with moderate fan pressure ratios. It is desirable to use an engine core that includes a low pressure turbine (LPT). Boost pressure and temperature to the high pressure compressor (HPC) is typically much higher in the low bypass derivative engine than in the original high bypass engine. This generally means that the maximum working air flow in the core is less than its total design modified air flow capacity due to the core physical maximum core rotational speed and / or mechanical limits of maximum compressor discharge temperature performance. Need to limit. It would be desirable to find a way to maximize the thrustable capacity of the derived engine by greatly increasing the fan pressure ratio to the bypass stream while operating the original engine core airflow at its maximum possible capacity.

  Thus, having a fan system that allows the original engine core to operate near its full airflow performance while greatly increasing the fan pressure ratio to the bypass stream to maximize the thrustable capacity of the derived engine. Is desirable.

  The above needs or needs can be met by an exemplary embodiment that provides a fan system, which includes a front fan stage configured to pressurize air flow, and an application from the front fan stage. And a rear fan stage with a leading fan configured to pressurize a first portion of the pressurized air stream, the rear fan stage being driven by the second portion of the pressurized air stream. ing.

  In one aspect of the invention, the rear fan stage rotates independently of the front fan stage.

  In another aspect of the invention, the aft fan stage includes a turbine airfoil adapted to extract energy from a pressurized air stream and a tip fan blade adapted to pressurize the air stream. Has a blade.

  The subject matter regarded as the invention is particularly pointed out and distinctly claimed in the claims appended hereto. However, the invention can best be understood by referring to the following detailed description taken in conjunction with the accompanying drawings.

1 is a schematic cross-sectional view of a portion of a gas turbine engine with an exemplary embodiment of an intermediate fan stage in accordance with the present invention. 1 is a schematic cross-sectional view of an exemplary gas turbine engine according to the present invention having an exemplary embodiment of an intermediate fan stage. FIG.

  Referring to the drawings wherein like reference numerals represent like elements throughout the various views, FIG. 1 illustrates an exemplary turbofan gas turbine engine 10 incorporating an exemplary embodiment of the present invention. The exemplary gas turbine engine 10 includes an engine center axis 11, a fan 12 that receives an inflow of ambient air 1, an optional booster or low pressure compressor (LPC) (not shown in FIG. 1), A high-pressure compressor (HPC) 18, a combustor 20 that mixes air and fuel pressurized by the HPC 18 to generate a combustion gas flowing downstream through a high-pressure turbine (HPT) 22, and thereby the combustion gas is engine 10 and a low-pressure turbine (LPT) 24 discharged from the engine 10. HPT 22 is coupled to HPC 18 using HPT shaft 23 to form a substantially high pressure rotor 29. The low pressure shaft 25 joins the LPT 24 to the fan 12 (and optional booster, if present) to form a substantially low pressure rotor 28. The second or low pressure shaft 25 is disposed coaxially with the high pressure rotor 29 and rotatably inside the high pressure rotor 29 in the radial direction. Although the low pressure rotor 28 and the high pressure rotor 29 are aerodynamically coupled, they rotate independently because they are not mechanically coupled.

  The HPC 18 that pressurizes the air flowing through the core has a rotor 19 that rotates about the longitudinal central axis 11. The HPC system includes a plurality of inlet guide vanes (IGV) 30 and a plurality of stator vanes 31 disposed circumferentially about the longitudinal central axis 11. The HPC 18 further includes a corresponding rotor blade 40 extending radially outward from a rotor hub 39 or a corresponding rotor in the form of a separate disk, or a plurality of integral drums, or an annular drum in any conventional manner. Rotor stage 19. The high pressure rotor 29 is supported within the engine stationary frame using known support methods using suitable bearings.

  Cooperating with each rotor stage 19 is a corresponding stator stage with a plurality of circumferentially spaced stator vanes 31. FIG. 1 shows an exemplary arrangement of stator vanes and rotor blades in an axial flow high-pressure compressor 18. The rotor blade 40 and the stator vane 31 form an airfoil having a corresponding aerodynamic profile or profile adapted to continuously pressurize the core air stream 8 in the axial stage. The rotor blade 40 rotates in an annular casing 38 surrounding the rotor blade tip. During operation, the pressure of the core air stream 8 increases as air decelerates and diffuses through the stator and rotor airfoils.

  FIG. 1 shows a fan system 50 that includes a front fan stage 52 that pressurizes the air stream 1. The pressurized air flow 2 flows out axially rearward from the front fan stage 52. A fixed annular splitter 46 which is coaxial with the central axis 11 is installed behind the front fan stage 52 in the axial direction. As shown in FIG. 1, the annular splitter 46 divides the pressurized air stream 2 into a first part 3 and a second part 4.

  The fan system 50 has a rear fan stage 60 that is installed axially behind the annular splitter 46. The rear fan stage 60 includes a rear fan rotor 61 and has a circumferential row of rear fan blades 62. The rear fan stage 60 rotates about the central axis 11 but is not mechanically coupled to the high pressure compressor 18 or the front fan stage 52. The rear fan stage 60 is aerodynamically coupled to the front fan stage 52 and the front stage of the high pressure compressor 18 during operation of the engine 10, but the rear fan stage 60 is mechanically coupled to the low pressure rotor 28 and the high pressure rotor 29. Rotate independently. Therefore, the rear fan stage 60 rotates independently of the front fan stage 52 installed upstream thereof.

  As shown in FIG. 1, the rear fan stage 60 includes a row of rear fan blades 62 disposed circumferentially about the longitudinal axis 11. Each rear fan blade 62 has a radially inner portion 63 and an outer portion 64. The radially inner portion 63 of the rear fan blade 62 is configured to be driven as an air turbine blade 82 that can extract energy from the pressurized air stream 7 flowing into the inner portion 63. Known air turbine airfoil shapes can be used to construct the inner portion 63 of the rear fan blade 62. As the airflow flows over the inner portion 63, the airflow expands to form an air outlet stream 57 having a lower pressure and lower temperature and energizes the rear fan blade 62 to cause the rear fan stage 60 to flow. To drive.

  As shown in FIG. 1, each rear fan blade 62 has an outer portion 64 and an arcuate shroud 65 disposed between the inner portion 63 and the outer portion 64. The outer portion 64 of the rear fan blade 62 is configured to be a leading fan blade 72 that can pressurize the air inflow 6. The arcuate shroud 65 supports the tip fan blade 72. The outer portion of the rear fan blade 62 has a known airfoil shape as a fan blade capable of pressurizing the air inflow 6. In the assembled state of the rear fan stage 60, the arcuate shroud 65 of each blade 62 abuts the arcuate shroud of the circumferentially adjacent fan blade 62 to form an annular platform and a tip fan 70 with a tip fan blade 72. Form. In one embodiment, each rear fan blade 62 has one tip fan blade 72 supported by an arcuate shroud 65. In another embodiment, each rear fan blade 62 can have a plurality of tip fan blades 72 supported by an arcuate shroud 65.

  As shown in FIG. 1, the rear fan stage 60 has a tip fan 70 configured to pressurize the first portion 3 of the pressurized air stream 2 from the front fan stage 52. The tip fan 70 is driven by a rear blade inner portion 63 that acts as an air turbine blade 82. The rear fan stage 60 with the tip fan 70 is driven by the second part 4 of the pressurized air flow 2. The inner portion 63 of the rear fan blade 62 is configured to act as an air turbine blade that can extract energy from the pressurized air stream, while the outer portion 64 of the rear fan blade 62 can pressurize the air flow. It is configured to be a compression type airfoil. Inner portion 63 is an air turbine blade 82 having a turbine-type airfoil 84 adapted to extract energy from a flow of pressurized air. Instead, the outer portion 64 of the rear fan blade 62 is referred to herein as the tip fan blade 72. The tip fan blade 72 can pressurize the air stream 6 to generate a pressurized tip flow 56 (see FIG. 1).

  As shown in FIG. 1, the fan system 50 further includes a circumferential row of inlet guide vanes (IGVs) 74 installed axially forward of the tip fan 70 of the rear fan stage 60. The IGV 74 has a known airfoil shape that can redirect the incoming air stream 3 to be the air stream 6 that flows into the tip fan 70 in a manner suitable for further pressurization. The IGV 74 is preferably supported by the inner casing 68 (see FIG. 1) and / or the splitter 46. To enhance control of the operation of the rear fan stage 60, the fan system 50 can have an inlet guide vane 74 with a variable vane configured to regulate the air flow 6 to the tip fan 70. The amount and orientation of the air flow 6 directed to the tip fan 70 can be changed by using a known actuator 75 to suitably move a portion of the IGV 74 to change the stagger angle.

  FIG. 2 illustrates an exemplary embodiment of a gas turbine engine 110 with a multi-stage fan 112 having a plurality of forward fan stages 152 configured to pressurize the air stream 1. Although the exemplary engine 110 shown in FIG. 2 shows three forward fan stages 152, any suitable number of forward fan stages can be selected for a particular application. The front fan stage pressurizes the flow stream 1 entering the fan to generate a pressurized flow stream 2. The forward fan stage is driven by a low pressure rotor 128 that includes a low pressure turbine 124 and a low pressure turbine shaft 125. The gas turbine engine 110 further includes a compressor 118 driven by a high pressure rotor 129 having a high pressure turbine 122 and a high pressure shaft 123. The HPC 118 has a rotor 19 that pressurizes air 8 that rotates about the central longitudinal axis 11 and flows through the core. The HPC system includes a plurality of stator vanes disposed circumferentially about a longitudinal central axis 11 (see, eg, FIG. 1). The HPC 118 further includes a corresponding rotor blade 140 extending radially outward from a corresponding rotor in the form of a rotor hub 139 or a separate disk, or a plurality of integral drums, or an annular drum in any conventional manner. Rotor stage 119. The high pressure rotor 129 is supported within the engine stationary frame using known support methods using suitable bearings. High pressure turbine 122 and low pressure turbine 124 are driven by combustion gases that are generated in combustor 120 and exit as hot exhaust stream 92.

  This exemplary embodiment of the gas turbine engine 110 includes an annular splitter 146 (see FIG. 2) located axially rearward of the axial final forward fan stage 152. The splitter 146 is adapted to branch the pressurized flow stream 2 from the front fan stage 152 to form the first portion 3 and the second portion 4 of the pressurized flow 2.

  As shown in FIG. 2, this exemplary embodiment of gas turbine engine 110 includes a rear fan stage 160 installed axially rearward of splitter 146 and axially forward of compressor 118. As shown in FIG. 2, the rear fan stage 160 has a tip fan 170 configured to pressurize the first portion 3 of the pressurized air stream 2 from the front fan stage 152. The tip fan 170 is driven by a rear blade inner portion 163 that acts as an air turbine blade 182. The rear fan stage 160 with the front fan 170 is driven by the second part 4 of the pressurized air flow 2. The inner portion 163 of the rear fan blade 162 is configured to act as an air turbine blade that can extract energy from the pressurized air stream, while the outer portion 164 of the rear fan blade 162 can pressurize the air flow. It is configured to be a compression type airfoil. Inner portion 163 is an air turbine blade 182 having a turbine-type airfoil 184 adapted to extract energy from a flow of pressurized air. Instead, the outer portion of the rear fan blade 162 is referred to herein as the leading fan blade 172. The tip fan blade 172 can pressurize the air stream 6 to generate a pressurized tip flow 56 (see, eg, FIG. 1). The rear fan stage 160 reduces the pressure and temperature of the pressurized air flow that drives the rear fan stage 160. Known air turbine airfoil shapes, materials and manufacturing methods can be used to construct the inner portion 163 of the rear fan blade 162. As the air flow flows over the inner portion 163, the air flow expands to form an air outflow 57 having a lower pressure and lower temperature and energizes the rear fan blade 162 to cause the rear fan stage 160 to flow. To drive.

  The exemplary gas turbine engine 110 shown in FIG. 2 further includes a circumferential row of inlet guide vanes (IGVs) 174 installed axially forward of the tip fan blades 172. Known airfoil shapes, materials, and manufacturing methods can be used to construct IGV174. The IGV 174 controls the volume of air flow into the tip fan 170 in the same manner as the apparatus shown in FIG. To enhance control of air flow into the tip fan 170, the inlet guide vane 174 is a variable vane configured to regulate the air flow to the tip fan 70. The amount and orientation of the air flow directed to the tip fan 710 can be changed by changing the stagger angle by suitably moving a portion of the IGV 174 using a known actuator 175.

  In one aspect of the invention, the exemplary gas turbine engine 110 shown in FIG. 2 (and FIG. 1) further flows an annular inner bypass passage 142 adapted to flow the inner bypass flow 56 and the outer bypass flow 5. And an annular outer bypass passage 144. The outer bypass flow 5 flows through the outer bypass passage 144 and is not pressurized by the tip fan 170. The inner bypass flow 6 (see FIG. 1) is pressurized by the tip fan 170 and flows out as a pressurized tip flow 56. A front mixer 148 located downstream of the rear fan stage 160 mixes the higher pressure inner bypass flow 56 and the lower pressure outer bypass flow 5 to form a mixed bypass flow 9 and generates a static pressure equilibrium. It is provided to strengthen it. As the mixer 148, a known mixer (alternatively known in the art as a variable area bypass injector or VABI) can be used. Backflow in the outer bypass passage 144 can be prevented by using a known blocker door 145 installed near the front region of the outer bypass passage 144. When the engine is operated, when the variable IGV 174 is opened and further pressurized by the front fan 170, the blocker door is operated in the closing direction. The gas turbine engine 110 further includes a rear mixer 94 (instead of being located downstream of the low pressure turbine 24 and adapted to enhance mixing of the hot exhaust 92 and the relatively cooler bypass airflow stream 91 from the low pressure turbine 24. , Known in the art as a variable area bypass injector or VABI). A known mixer (VABI) can be used for this purpose. During engine operation, the operability of the front fan stage 152 and the rear fan stage 160 should be suitably scheduled using known methods, i.e., the actions of the variable IGV 144, blocker door 145, front mixer 148 and rear VABI 194. Can be controlled by.

  As shown in FIGS. 1 and 2, the rear fan stages 60, 160 (alternatively referred to herein as medium pressure fan stages or IPFS) are core drive fans coupled to a core spool as is known in the art. Unlike the stage, it is a separate independent rotating spool incorporating the front fan 70, 170. In addition, as described herein, the IPFS has tip fan blades 72, 172 in its outer portion and air turbine blades 82, 182 in its inner portion. The IPFS spool is installed between the front fans 52 and 152 and the HPCs 18 and 118 so that a part of the fan air is supplied to the tip of the IPFS, and the pressure at the tip of the IPFS is the IPFS tip fan blade 72, 172 and then fed to the inner bypass passages 42, 142. The inner part 4 of the front fan stream 2 is fed to the turbine blades 82, 182 in the IPFS inner part, where the inner part of the front fan stream 2 expands to drive the fan tip. To give. The flow from the turbine outlet is fed to the inlets of HPCs 18,118. The extraction of energy by the IPFS turbine blades 82, 182 reduces the boost pressure and temperature into the HPC below the boost pressure and temperature at the front fan outlets 52, 152. By judicious selection of the pressure ratio of the front fans 52, 152 and the IPFS tip fans 70, 170, the inlet conditions for the high pressure compressors 18, 118 are matched to the original (reference) engine design requirements and the core flow capability by the derived engine. The use of can be maximized. At the same time, the front fans 52, 152 and the IPFS 60, 160 provide the desired higher bypass air pressure for the bypass flow 9.

  Repeated investigations have shown that the possible thrust in existing cores can be increased by up to 20% over derived mixed flow turbofans at the same fan airflow dimensions. The temperature level to the HPC can be easily matched to the original hardware design conditions that allow maximum use of the correction flow capability within the original core machine design limits. One skilled in the art will appreciate that a flow path architecture study using known methods can be performed to set the required mounting structure for IPFS and the aerodynamic design characteristics of the fan tip and turbine hub. . In this exemplary embodiment shown herein, the IPFS is preferably mounted within the fan frame structure and thus does not require an additional main engine frame to mount an additional pool.

  With reference to FIGS. 1 and 2, an exemplary method of operating fan systems 50, 150 includes the following steps. The air stream 1 flowing in the fan systems 50, 150 generates a pressurized stream 2 that is pressurized in the front fan stages 52, 152 and flows out of the front fan stages. The pressurized stream 2 is branched into the first part 3 and the second part 4 using suitable means, for example using annular splitters 46, 146. The first portion 3 of the pressurized air stream 2 is then directed towards the tip fan 70 of the rear fan stage 60. A portion of the pressurized flow 2 flows through the outer bypass passages 44, 144 to generate the outer bypass flow 5. The rear fan stage 60 rotates independently of the front fan stage 52. The second portion 4 of the pressurized air stream 2 is directed towards the circumferential row of air turbine blades 82 of the rear fan stage 60 so that the rear fan stage 60 is driven by the pressurized air. . At this point, the higher pressure inflow 7 that flows into the inner portions 63, 163 of the rear fan stage 60 expands to a lower pressure outflow 57. During this expansion, the temperature of the expanded air flow in the inner portions 63 and 163 of the rear stages 60 and 160 decreases. Accordingly, the temperature and pressure of the core stream 8 flowing into the compressors 18 and 118 are reduced.

  This exemplary method further includes pressurizing the stream 6 entering the tip fans 70, 170 to generate a pressurized tip stream 56 (see FIGS. 1 and 2). The air flow 6 flowing into the tip fan 70 is adjusted by the inlet guide vanes 74 and 174. Specifically, the amount of air flowing through the tip fans 70, 170 is controlled independently by the inlet guide vanes 74, 174. More specifically, the staggers of the inlet guide vanes 74, 174 are configured to selectively control the amount of air flow through the tip fans 70, 170 based on the fan pressure ratio, thrust and performance requirements of the engines 10, 110. Changed to The adjustment of the air 6 between the nearly zero air flow rate and the maximum discharge air flow rate is performed by changing the stagger of the inlet guide vanes 74 and 174 as necessary. In this exemplary embodiment, inlet guide vanes 74, 174 are mechanically operated by known actuators 75, 175 and actuated by a known main engine control system (not shown). In another embodiment, the inlet guide vanes 74, 174 are actuated by any suitable mechanism. In addition, the exemplary method mixes the outer bypass flow 5 in the annular outer bypass passages 44, 144 with the tip flow 56 from the tip fans 70, 170 in the annular inner bypass passages 42, 142 to provide a mixed bypass flow. 9 is formed. Doors 45, 145 located near the outer bypass passages 44, 144 are positioned within the outer bypass passage 144 by adjusting the blocker doors 45, 145 between a partially closed position and a substantially fully open position. It works to prevent backflow in A mechanical actuator operated by a known main engine control system (not shown) is used in the exemplary embodiments shown herein. The method described herein optionally includes the step of operating a known type of forward mixer to control the mixing of the outer bypass flow 5 and the tip flow 56 and to achieve a suitable hydrostatic equilibrium. Further, the method includes the step of operating a known type of rear mixer 94, 194 to control the operating characteristics of the front fan stage 152 and the rear fan stage 160 and the performance of the engine 10, 110. The front mixers 48, 148, the rear mixers 94, 194, the blocker doors 45, 145, and the inlet guide vanes 74, 174 are operated in a controlled manner using an engine control system (not shown) to operate the engines 10, 110. Try to maximize properties and performance.

  This written description uses examples, including the best mode, to disclose the invention and also to enable any person skilled in the art to make and use the invention. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other embodiments may have structural elements that do not differ from the language of the claims, or they contain equivalent structural elements that have non-essential differences from the language of the claims. Is intended to fall within the scope of the appended claims.

DESCRIPTION OF SYMBOLS 1 Ambient air 2 Pressurized air flow from front fan stage 3 First part of pressurized air flow from front fan stage 4 Second part of pressurized air flow from front fan stage 5 Outer bypass flow 6 Inner bypass Stream 7 Pressurized air stream to the rear fan stage 8 Core air stream 9 Mixed bypass stream 10, 110 Gas turbine engine 11 Engine center axis 12, 112 Fan 18, 118 High pressure compressor (HPC)
19, 119 Rotor stage 20, 120 Combustor 22, 122 High pressure turbine (HPT)
23, 123 HPT shaft 24, 124 Low pressure turbine (LPT)
25, 125 Low pressure shaft 28, 128 Low pressure rotor 29, 129 High pressure rotor 30 Inlet guide vane (IGV)
31 Stator vane 38 Casing 39, 139 Rotor hub 40, 140 Rotor blade 42, 142 Annular inner bypass passage 44, 144 Annular outer bypass passage 45, 145 Blocker door 46, 146 Splitter 48, 148 Forward mixer 50 Fan system 52, 152 Forward fan stage 56 Pressure front flow 57 Outflow flow from rear fan blades 60, 160 Rear fan stage 61 Rear fan rotors 62, 162 Rear fan blades 63, 163 Radial inner portions 64, 164 of rear fan blades radially outward Portion 65 Arcuate shroud 68 Inner casing 70, 170 Tip fan 72, 172 Tip fan blade 74, 174 Inlet guide vane (IGV)
75, 175 Actuators 82, 182 Air fan blades 84, 184 Turbine type airfoil 91 Cold bypass air flow stream 92 Hot exhaust stream 94, 194 Rear mixer

Claims (12)

A gas turbine engine (10) comprising:
A mechanically driven forward fan stage (52) configured to pressurize the air stream (1);
A compressor (18);
A first portion (2) of the pressurized air flow (2) from the front fan stage (52) is installed axially rearward and axially forward from the compressor (18). 3) a rear fan stage (60) with a circumferential row of tip fan blades (72) adapted to pressurize,
The rear fan stage (60) is driven by energy extracted from the second part (4) of the pressurized air stream (2);
The rear fan stage (60) rotates independently of the front fan stage (52) and the compressor (18);
Gas turbine engine. The gas turbine engine of claim 1, wherein the rear fan stage (60) reduces the pressure in the second portion (4) of the pressurized air flow (2) from the front fan stage (52). The gas turbine engine according to claim 1 or 2, further comprising a circumferential row of inlet guide vanes (74) installed axially forward from the tip fan blades ( 72 ). The gas turbine engine of claim 3, wherein the inlet guide vane (74) is a variable vane configured to regulate an air flow (6) to the tip fan blade ( 72 ). A first portion (3) of the pressurized air stream (2) flows upstream of the tip fan blade (72), flows through an annular inner bypass passage (142) and is pressurized by the tip fan blade (72). The inner bypass flow (56) and the outer bypass flow (5) that flows through the annular outer bypass passage (144) provided radially outside the annular inner bypass passage (142) and is not pressurized by the tip fan blade (72). And divided into,
The gas turbine engine according to any one of claims 1 to 4.
The gas turbine engine of claim 5, further comprising a blocker door (145) that prevents backflow in the outer bypass passage (144). An annular splitter (46) installed in front of the rear fan stage (60) in the axial direction and dividing the fluid flow from the front fan stage into a first part (3) and a second part (4) is further provided. The gas turbine engine according to any one of claims 1 to 6, further comprising: Installed downstream of the rear fan stage (160);
The gas turbine engine of claim 5, further comprising a forward mixer (148) that promotes mixing of the inner bypass flow (56) and the outer bypass flow (5) to form a mixed bypass flow (9). The rear mixer (94) installed downstream of the low pressure turbine (24) to enhance mixing of hot exhaust (92) and a relatively cooler air stream from the low pressure turbine (24). The gas turbine engine according to any one of the above. A method of operating a gas turbine engine (110) comprising:
Pressurizing the air stream (1) using a mechanically driven forward fan stage (152) to generate a pressurized air stream (2);
Directing the first portion (3) of the pressurized air stream (2) toward the leading fan (170) of the rear fan stage (160);
What is said front fan stage (152) by expanding a second part (4) of said pressurized air stream (2) in a circumferential row of air turbine blades (182) of said rear fan stage (160)? Driving the independently rotating rear fan stage (160) to reduce the pressure of the core flow (5) flowing into the compressor (118);
including,
Method. The method of claim 10, further comprising adjusting an air flow (6) entering the tip fan (170) with an inlet guide vane (174). 12. The method of claim 11, further comprising adjusting the air flow (6) between a substantially zero air flow rate and a maximum discharge air flow rate.

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