JP5454083B2 - Compressor for jet engine and jet engine - Google Patents

Description

  The present invention relates to a compressor for a jet engine that compresses air taken in an annular core flow path formed in a cylindrical engine case in a jet engine.

  A conventional jet engine compressor will be described with reference to FIGS. 9 and 10. FIG. In the figure, “F” indicates the forward direction (upstream direction), and “R” indicates the backward direction (downstream direction).

  As shown in FIG. 9, a jet engine compressor 101 according to the prior art includes a cylindrical compressor case 103, which is one of the cylindrical engine cases 105 in the jet engine. Part. A compressor rotor 107 having a plurality of stages (only the final stage is shown) is provided in the compressor case 103 so as to be rotatable along the axial direction (the axial direction of the compressor case 103). The machine rotor 107 includes a disk 109 that can rotate around an axis (the axis of the compressor case 103), and a core flow path 111 that is provided on the outer peripheral surface of the disk 109 at equal intervals and is formed in the engine case 105. A plurality of (only one shown) moving blades 113 are provided. Further, in the compressor case 103, a plurality of stages (only the last stage is shown) of compressor stators 115 are provided alternately with a plurality of stages of compressor rotors 107 along the axial direction. 115 includes a plurality (only one shown) of stationary blades 117 disposed at equal intervals in the circumferential direction and positioned in the core channel 111.

  Between the compressor stator 115 at the final stage and the combustor 119 in the jet engine, an annular diffuser 121 that decelerates compressed air (compressed air) and sends it to the combustor 119 is formed. A part of the core channel 111 is configured. A plurality of struts 123 that support the compressor case 103 are provided in the diffuser 121 at intervals in the circumferential direction.

  Therefore, the compressor rotor 107 for jet engine is driven to rotate the multi-stage compressor rotor 107, so that the multi-stage compressor rotor 107 and the multi-stage compressor stator 115 cooperate with each other into the core flow path 111. The air can be compressed. Then, the compressed air is decelerated by the diffuser 121 and sent to the combustor 119.

  By the way, in recent years, various developments have been made in order to reduce the weight of the jet engine, and as a part thereof, a jet engine compressor 125 in which the strut 123 is omitted as shown in FIG. 10 has also been developed. Specifically, the final stage compression is performed so that the plurality of stationary blades (final stage stationary blades) 117 of the final stage compressor stator 115 have a function of supporting the compressor case 103 (function as a strut). Each stator blade (the last stage stator blade) 117 of the machine stator 115 is thicker than the stator blades (each other stage stator blade) 117 of the compressor stator 117 of the other stage. The length is getting longer. Accordingly, the length from the inlet of the compressor stator 115 at the final stage to the outlet of the diffuser 121 is shortened compared to the jet engine compressor 101 according to the related art having the struts 123, in other words, the jet The engine shaft length can be shortened to reduce the weight of the jet engine.

  In addition, there exist some which are shown to patent document 1 and patent document 2 as a prior art relevant to this invention.

JP 2003-13748 A Japanese Patent Application Laid-Open No. 08-61093

  However, the jet engine compressor 125 according to the related art, in which the struts are omitted, can reduce the length of the jet engine by reducing the length from the inlet of the final stage compressor stator 117 to the outlet of the diffuser 121. Although it is possible, the compressed air is rapidly decelerated between the inlet of the final stage compressor stator 117 and the outlet of the diffuser by an amount corresponding to the shortened length from the inlet of the final stage compressor stator 117 to the outlet of the diffuser 121. It is necessary to let Therefore, as shown in FIGS. 11 (a) and 11 (b), a region where energy loss due to separation of the flow of the compressed air is large in the vicinity of the blade surface of the final stage stationary blade 117 and the wall surface of the diffuser 121 (the energy loss is extremely low). A large area), causing a reduction in the compression efficiency of the jet engine compressor 125, and the presence of a large energy loss area (low energy fluid) in the vicinity of the wall surface of the diffuser 121. There is a problem that the direction changes with respect to the direction required by the combustor 119 and the combustion efficiency of the combustor 119 is reduced. In addition, the area | region with a large energy loss in Fig.11 (a) (b) is calculated | required by three-dimensional steady viscosity CFD (Computational Fluid Dynamics) analysis.

  Accordingly, an object of the present invention is to provide a jet engine compressor having a novel configuration that can solve the above-described problems.

  As a result of repeating trial and error in order to solve the above-mentioned problems, the inventor of the present invention, as shown in FIG. 2, compresses the final stage on the radially inner wall surface (wall surface on the hub side) of the core channel. The part between the leading edge side and the trailing edge side of the stator blade (final stage stationary blade) of the machine stator (predetermined part, more specifically, the part from the leading edge to the trailing edge of the final stage stationary blade) is appropriate When a predetermined part on the radially inner wall surface of the core channel is a straight line when the concave first curved surface is changed to a convex second curved surface with reference to an inflection point (shown in FIG. 10) 3 (a) and (b), as shown in FIGS. 3 (a) and 3 (b), the energy loss region on the blade surface of the stationary blade and the wall surface of the diffuser is large. New knowledge that it can be reduced, and the present invention is completed. It led to. This increases the deceleration of the compressed air near the inlet of the final stage compressor stator in the core flow path, resulting in the compressed air near the outlet of the final stage compressor stator in the core flow path. This is thought to be due to slow deceleration. Here, the appropriate inflection point is a point that is 70 to 85% cord length, preferably 75 to 80% cord length away from the leading edge of the stator blade of the final stage compressor stator. In addition, the area | region with large energy loss in Fig.3 (a) (b) is calculated | required by the three-dimensional steady viscosity CFD analysis.

  A first feature of the present invention is a jet engine compressor that compresses air taken into an annular core flow path formed in a cylindrical engine case of a jet engine, and constitutes a part of the engine case A cylindrical compressor case that is configured to be rotatable along an axial direction (axial direction of the compressor case) in the compressor case and rotatable about an axis (axial center of the compressor case) And a plurality of compressor rotors provided with a plurality of rotor blades integrally provided at equal intervals on the outer peripheral surface of the disk and positioned in the core flow path, and an axial direction in the compressor case A plurality of compressor stators provided alternately with a plurality of compressor rotors along the circumferential direction, arranged at equal intervals in the circumferential direction and provided with a plurality of stationary blades located in the core flow path, And the pressure in the final stage Each stator blade in the compressor stator at the final stage has each stator blade of the compressor stator at the other stage so that each stator blade of the machine stator has a function of supporting the compressor case (function as a strut). The cord length is longer than that of the blades, and a part of the core flow path is formed between the compressor stator at the final stage and the combustor in the jet engine, and the compressed air is decelerated and the combustion is performed. An annular diffuser that is fed to the compressor is formed, and a portion from the leading edge to the trailing edge of the stator blade of the compressor stator at the final stage on the radially inner wall surface (hub-side wall surface) of the core channel The (predetermined part) transitions from the concave first curved surface to the convex second curved surface, and the inflection point transitioning from the first curved surface to the second curved surface is the compressor stator at the final stage. From the leading edge of the stationary blade And summarized in that a point at a distance of 0 to 85% chord length.

  According to the first feature, the core is formed by the cooperation of a plurality of stages of the compressor rotor and a plurality of stages of the compressor stator by rotating the plurality of stages of the compressor rotor by driving the jet engine compressor. The air taken into the flow path can be compressed. Then, the compressed air is decelerated by the diffuser and sent to the combustor.

  Further, since each stator blade of the compressor stator at the final stage has a function of supporting the compressor case, struts can be omitted from the jet engine compressor. Thereby, compared with the compressor for jet engines equipped with the strut (see FIG. 9), the length from the compressor stator inlet of the final stage to the outlet of the diffuser is shortened, in other words, It is possible to reduce the weight of the jet engine by shortening the axial length of the jet engine.

  And the predetermined site | part in the wall surface of the radial inside of the said core flow path has changed from the said 1st curved surface of concave shape to the said 2nd curved surface of convex shape, and changes from the said 1st curved surface to the said 2nd curved surface Since the inflection point is a point that is 70 to 85% cord length away from the leading edge of the stationary blade of the compressor stator at the final stage, applying the above-described novel knowledge, the radially inner side of the core channel Compared with the case where the predetermined part of the wall surface of the slab is a straight line, it is possible to reduce the region of large energy loss on the blade surface of the stationary blade and the wall surface of the diffuser.

  According to a second aspect of the present invention, a jet engine includes the jet engine compressor having the first characteristic.

  According to the 2nd characteristic, there exists an effect | action similar to the effect | action by a 1st characteristic.

  According to the present invention, the axial length of the jet engine is shortened to reduce the weight of the jet engine, and compared with a case where a predetermined portion on the radially inner wall surface of the core channel is a straight line. Thus, it is possible to reduce a region of large energy loss on the blade surface of the stationary blade and the wall surface of the diffuser, so that the compression efficiency of the compressor for the jet engine is improved and the direction of the flow of the compressed air is increased. A change in the direction required by the combustor can be suppressed, and the combustion efficiency of the combustor can be increased.

It is a typical side view showing the important section of the high-pressure compressor for jet engines concerning the embodiment of the present invention. It is a figure explaining the novel knowledge which the inventor of this invention discovered. Fig. 3 (a) is a schematic view from the meridional plane showing the region near the suction surface of the final stage stationary blade and the energy loss in the diffuser, and Fig. 3 (b) is a new finding. It is a typical figure from the IIIB arrow of FIG. 3 (a) which shows the area | region where the energy loss is large in the exit vicinity of the diffuser which concerns. It is a side view of the jet engine which concerns on embodiment of this invention, Comprising: The upper half is cross-sectioned. It is a typical side view which shows the periphery of the stationary blade of the last stage which concerns on the modification 1 of embodiment of this invention. FIG. 6A is a schematic view from the meridional surface direction showing the vicinity of the suction surface of the final stage stationary blade according to the first modification of the embodiment of the present invention and the region where the energy loss in the diffuser is large, and FIG. FIG. 6B is a schematic view from the VIB arrow of FIG. 6A showing a region with a large energy loss in the vicinity of the exit of the diffuser according to the first modification of the embodiment of the present invention. FIG. 7A is a view showing a blade cross section on the tip side of the final stage stationary blade according to the second modification of the embodiment of the present invention, and FIG. 7B is a second modification of the embodiment of the present invention. It is a figure which shows the blade | wing cross section of the center span side of the stationary blade of the last stage which concerns. FIG. 8A is a schematic diagram from the meridional plane direction showing the vicinity of the suction surface of the final stage stationary blade according to the second modification of the embodiment of the present invention and the region where the energy loss in the diffuser is large, and FIG. FIG. 8B is a schematic view from the direction of arrow VIIIB in FIG. 8A, showing a region with a large energy loss in the vicinity of the exit of the diffuser according to the second modification of the embodiment of the present invention. It is a typical side view which shows the compressor for jet engines which concerns on a prior art. It is a typical side view which shows the compressor for jet engines which concerns on the prior art which abbreviate | omitted the strut. FIG. 11A is a schematic view from the meridional plane direction showing the vicinity of the suction surface of the final stage stationary blade and the large energy loss region in the diffuser according to the prior art, and FIG. It is a typical figure from the XIB arrow view of Fig.11 (a) which shows the area | region where an energy loss is large in the exit vicinity of the diffuser concerned.

  An embodiment of the present invention will be described with reference to FIGS. In the figure, “F” indicates the forward direction (upstream direction), and “R” indicates the backward direction (downstream direction).

  As shown in FIG. 4, the jet engine 1 according to the embodiment of the present invention is an engine mounted on an aircraft, and includes a cylindrical engine case 3 as a base. A cylindrical cowl 5 is provided so as to surround it. An annular core channel (main channel) 7 is formed in the engine case 3, and an annular bypass channel 9 is formed between the cowl 5 and the engine case 3.

  A fan (fan rotor) 11 for taking air into the core flow path 7 and the bypass flow path 9 is rotatably provided at the front portion of the engine case 3. A low-pressure compressor 13 is provided on the rear side of the fan 11 in the engine case 3 to compress the air taken into the main flow path 7 at a low pressure. The low-pressure compressor 13 includes a plurality of low-pressure compressor rotors 15 provided in the engine case 3 so as to be rotatable in the axial direction (axial direction of the engine case 3), and the engine case 3 in the axial direction. A plurality of low-pressure compressor rotors 17 provided alternately with a plurality of low-pressure compressor rotors 15 are provided.

  A high-pressure compressor 19 is provided on the rear side of the low-pressure compressor 13 in the engine case 3 to compress the compressed air compressed at a low pressure. The high-pressure compressor 19 includes a plurality of high-pressure compressor rotors 21 provided in the engine case 3 so as to be rotatable in the axial direction, and a plurality of high-pressure compressors in the engine case 3 along the axial direction. A plurality of high-pressure compressor stators 23 provided alternately on the rotor 21 are provided. The details of the configuration of the high-pressure compressor 19 will be described later.

  An annular combustor 25 that burns fuel in compressed air is provided on the rear side of the high-pressure compressor 19 in the engine case 3. The combustor 25 includes a plurality of (only one shown) injection nozzles 27 that are arranged in the circumferential direction and inject fuel.

  A high-pressure turbine 29 is provided on the rear side of the combustor 25 in the engine case 3, and the high-pressure turbine 29 is driven by the expansion of combustion gas from the combustor 25 and interlocks with the high-pressure compressor 19. It is to be driven. The high-pressure turbine 29 includes a plurality of (two-stage) high-pressure turbine rotors 31 provided in the engine case 3 so as to be rotatable in the axial direction, and a plurality of high-pressure turbines 31 in the engine case 3 along the axial direction. A high-pressure turbine stator 33 is provided so as to be sandwiched between the turbine rotors 31. Here, the multiple-stage high-pressure turbine rotor 31 is integrally connected to the multiple-stage high-pressure compressor rotor 21.

  A low-pressure turbine 35 is provided on the rear side of the high-pressure turbine 29 in the engine case 3, and the low-pressure turbine 35 is driven by the expansion of the combustion gas and simultaneously drives the low-pressure compressor 13 and the fan 11. Is. The low-pressure turbine 35 includes a plurality of low-pressure turbine rotors 37 provided in the engine case 3 so as to be rotatable in the axial direction, and a plurality of low-pressure turbine rotors 37 provided in the engine case 3 along the axial direction. A plurality of low-pressure turbine stators 39 provided alternately are provided. Here, the multi-stage low-pressure turbine rotor 37 is integrally connected to the multi-stage low-pressure compressor rotor 15 and the fan 11.

  Next, a general operation of the jet engine 1 will be described.

  The high-pressure compressor 19 is driven by the operation of an appropriate starter device (not shown) to rotate the multi-stage high-pressure compressor rotor 21. Thereby, the operation of the jet engine 1 can be started.

  After the operation of the jet engine 1 is started, the high-pressure turbine 29 and the low-pressure turbine 35 are driven by the expansion of the combustion gas by combusting the fuel in the compressed air by the combustor 25, so The low pressure turbine rotor 37 of the stage is rotated. In addition, the high-pressure turbine 29 is driven in conjunction with the high-pressure turbine 29 to rotate the high-pressure compressor rotor 21 in multiple stages, and the low-pressure turbine 35 is driven in conjunction with the low-pressure compressor 13 and the fan 11. The multi-stage low-pressure compressor rotor 15 and the fan 11 are rotated. As a result, the air is taken into the core flow path 7 and the bypass flow path 9 by the fan 11, the air taken into the core flow path 7 by the low pressure compressor 13 is compressed at a low pressure, and further compressed by the high pressure compressor 19. Compressed air can be compressed at high pressure.

  A series of operations as described above (drive of the fan 11, drive of the low pressure compressor 13, drive of the high pressure compressor 19, combustion by the combustor 25, drive of the high pressure turbine 29, drive of the low pressure turbine 35) are continuously performed. As a result, the jet engine 1 can be appropriately operated, and a propulsive force can be obtained by the combustion gas injected from the core flow path 7 and the air injected from the bypass flow path 9.

  Next, details of the high-pressure compressor (high-pressure compressor for jet engine) 19 that is a main part of the embodiment of the present invention will be described.

  As shown in FIG. 1, the high-pressure compressor 19 according to the embodiment of the present invention includes a cylindrical high-pressure compressor case 41, and the high-pressure compressor case 41 constitutes a part of the engine case 3. To do. In the high-pressure compressor case 41, as described above, the high-pressure compressor rotor 21 having a plurality of stages (only the final stage is shown in FIG. 1) is provided along the axial direction (the axial direction of the high-pressure compressor case 41). The high-pressure compressor rotor 21 of each stage is provided with a disk 43 that can rotate around an axis (axis of the high-pressure compressor case 41), and an outer circumferential surface of the disk 43 provided at equal intervals and a core. A plurality of moving blades 45 located in the flow path 7 are provided.

  In the high-pressure compressor case 41, as described above, the high-pressure compressor stator 23 having a plurality of stages (only the final stage is shown in FIG. 1) is provided alternately with the plurality of high-pressure compressor rotors 21 along the axial direction. Each stage of the high-pressure compressor stator 23 includes a plurality (only one is shown in FIG. 1) of stationary blades 47 that are arranged at equal intervals in the circumferential direction and are positioned in the core flow path 7. . The final stage high-pressure compressor is configured so that each stationary blade (final stage stationary blade) 47 of the final-stage high-pressure compressor stator 23 supports the high-pressure compressor case 41 (function as a strut). Each stator blade 47 in the stator 23 has a longer cord length than each stator blade (each other stator blade) 47 of the other stage high-pressure compressor stator 23. In addition, it is desirable that each stationary blade 47 in the high-pressure compressor stator 23 in the final stage is thicker than each stationary blade 47 in the other stages.

  An annular diffuser 49 that decelerates the compressed air and sends it to the combustor 25 is formed between the high-pressure compressor stator 23 in the final stage and the combustor 25, and this diffuser 49 is a part of the core flow path 7. It constitutes.

  A portion (predetermined portion) from the front edge to the rear edge of the stationary blade 47 of the high-pressure compressor stator 23 at the final stage on the radially inner wall surface (hub-side wall surface) of the core flow path 7 is a concave shape. The transition from the first curved surface FR to the convex second curved surface SR is made. The inflection point IP at which the first curved surface FR transitions to the second curved surface SR is 70 to 85% code length from the leading edge of the stationary blade 47 of the high-pressure compressor stator 23 in the final stage, preferably 75 to 80%. This is the point separated by the length of the code. Here, if the inflection point IP is separated from the front edge of the stationary blade 47 of the high-pressure compressor stator 23 in the final stage by 70% or more, the core flow This is because it is difficult to increase the deceleration of the compressed air in the vicinity of the inlet of the final high-pressure compressor stator 23 in the path 7. On the other hand, the inflection point IP is not separated from the leading edge of the stationary blade 47 of the high-pressure compressor stator 23 in the final stage by exceeding 85% cord length. This is because it is difficult to moderately slow down the compressed air in the vicinity of the outlet of the final stage high-pressure compressor stator 23 in the path 7. The predetermined portion is not limited to the portion from the leading edge to the trailing edge of the stationary blade 47, and may be a portion between the leading edge side and the trailing edge side of the stationary blade 47.

  Then, the effect | action and effect of embodiment of this invention are demonstrated.

  As described above, by rotating the multi-stage high-pressure compressor rotor 21 by driving the high-pressure compressor 19, the core flow path is obtained by the cooperation of the multi-stage high-pressure compressor rotor 21 and the multi-stage high-pressure compressor stator 23. The air taken in 7 can be compressed at high pressure. Then, the compressed air is decelerated by the diffuser 49 and sent to the combustor 25.

  In addition, since each stationary blade 47 of the high-pressure compressor stator 23 at the final stage has a function of supporting the high-pressure compressor case 41, struts can be omitted from the high-pressure compressor 19. This shortens the length from the inlet of the high-pressure compressor stator 23 at the final stage to the outlet of the diffuser 49, in other words, compared to a jet engine compressor (see FIG. 9) having struts. The axial length of the engine 1 can be shortened to reduce the weight of the jet engine 1.

  And the predetermined site | part in the wall surface inside radial direction of the core flow path 7 has changed from the 1st curved surface FR to the convex 2nd curved surface SR from the concave 1st curved surface FR, and changes from the 1st curved surface FR to the 2nd curved surface SR. Since the inflection point IP is a point separated by 70 to 85% code length from the leading edge of the stationary blade 47 of the high-pressure compressor stator 23 in the final stage, the analysis result applying the above-described novel knowledge (FIG. 3 (a ) (See (b)) and the analysis result (see FIGS. 11 (a) and 11 (b)) when the predetermined portion of the wall on the radially inner side of the core channel 7 is a straight line (see FIG. 10) (see FIGS. 11 (a) and 11 (b)). It has been found that a region with a large energy loss (including a region with a very large energy loss) on the blade surface of the blade 47 and the wall surface of the diffuser 49 is reduced.

  Therefore, according to the embodiment of the present invention, the axial length of the jet engine 1 is shortened to reduce the weight of the jet engine 1, and the predetermined portion on the radially inner wall surface of the core flow path 7 is a straight line. Compared with the case of the above, since it is possible to reduce the large energy loss region on the blade surface of the stationary blade 47 and the wall surface of the diffuser 49, the compression efficiency of the high pressure compressor 19 is improved and the direction of the flow of compressed air Therefore, the combustion efficiency of the combustor 25 can be increased by suppressing the change in the direction required by the combustor 25.

(Modification 1)
A first modification of the embodiment of the present invention will be described with reference to FIGS. 5 and 6A and 6B.

As shown in FIG. 5, corresponding to the direction along the channel center PC of the diffuser 49, specifically, gradually from the axis of the high pressure compressor case 41 channel center PC of the diffuser 49 toward the downstream side in response to that is way away, the front edge 47 (stationary blade in the final stage) vanes of the high pressure compressor stator 23 of the last stage, the hub-side (base end side) chip side (distal side ) Is swept so that it is located on the downstream side.

  Here, the incident angle θh on the hub side of the leading edge of the final stage stationary blade 47 is 9 °, and the incident angle θt on the tip side of the leading edge of the stationary blade 47 of the final stage is 15 °. Further, the portion from the hub side of the leading edge of the stationary blade 47 of the final stage to the 60% span length is a straight line, and from the 60% span length of the leading edge of the stationary blade 47 of the final stage to the tip side. The part is a quadratic curve.

According to the first modification of the embodiment of the present invention, since in correspondence with the direction along the channel center PC of the diffuser 49 is made to sweep the leading edge of the vanes 47 of the last stage, before the vanes 47 of the last stage Compared to the analysis result (see FIGS. 3A and 3B) when the edge is not swept (as shown in FIG. 1), as shown in FIGS. 6A and 6B, the energy loss on the wall surface of the diffuser 49 It has been found that a large area (including a very large area of energy loss) can be further reduced. In addition, the area | region where an energy loss is large in FIG.

  Therefore, according to the first modification of the embodiment of the present invention, the effect of the above-described embodiment of the present invention can be further enhanced.

When the flow path center PC of the diffuser 49 is gradually approaching the axial center of the high-pressure compressor case 41 toward the downstream side, the front edge of the stationary blade 47 at the final stage is connected to the tip side on the hub side. It is desirable to sweep so as to be located on the upstream side.

(Modification 2)
A second modification of the embodiment of the present invention will be described with reference to FIGS. 7 and 8A and 8B.

As shown in FIG. 7, similarly to the first modification of the embodiment of the present invention, the leading edge of the vanes 47 of the last stage on which is swept, in correspondence with the direction along the channel center PC of the diffuser 49 Specifically, corresponding to the fact that the flow path center PC of the diffuser 49 gradually moves away from the axis of the high-pressure compressor case 41 toward the downstream side, The corner gradually decreases from the center span side to the tip side. The camber angle of the final stage stationary blade 47 is determined by (inlet swirl angle θg) − (exit swirl angle θe).

According to the second modification of the embodiment of the present invention, after in response to the direction along the channel center PC of the diffuser 49 is swept leading edges of the vanes 47 of the last stage, the channel center PC of the diffuser 49 Corresponding to the direction along which the camber angle of the final stage stationary blade 47 gradually decreases from the center span side to the tip side, the leading edge of the final stage stationary blade 47 is not swept (FIG. 1). As shown in FIGS. 8 (a) and 8 (b), compared with the analysis results (see FIGS. 3 (a) and 3 (b)), a region where the energy loss is large on the wall surface of the diffuser 49 (the energy loss is extremely low). (Including a large area) can be further reduced. In addition, the area | region with large energy loss in Fig.8 (a) (b) is calculated | required by the three-dimensional steady viscosity CFD analysis.

  Therefore, according to the second modification of the embodiment of the present invention, the effect of the above-described embodiment of the present invention can be further enhanced.

When the flow path center PC of the diffuser 49 is gradually approaching the axial center of the high-pressure compressor case 41 toward the downstream side, the camber angle of the final stage stationary blade 47 is increased from the center span side. It is desirable to make it gradually decrease toward the hub side.

Also, not by sweeping the front edge of the channel center vane of the last stage in response to the direction along the PC 47 of the diffuser 49, in response to the direction along the channel center PC of the diffuser 49, the final stage stationary blade Even when the camber angle of 47 is gradually reduced from the center span side to the chip side, an analysis result similar to the analysis result shown in FIGS. 8A and 8B is obtained, although illustration is omitted. It has been found.

  In addition, this invention is not restricted to description of the above-mentioned embodiment, It can implement in a various aspect. Further, the scope of rights encompassed by the present invention is not limited to these embodiments.

DESCRIPTION OF SYMBOLS 1 Jet engine 3 Engine case 5 Cowl 7 Core flow path 7 Main flow path 9 Bypass flow path 11 Fan 13 Low pressure compressor 19 High pressure compressor 21 High pressure compressor rotor 23 High pressure compressor stator 25 Combustor 29 High pressure turbine 35 Low pressure turbine 41 High pressure Compressor case 43 Disk 45 Moving blade 47 Stator blade 49 Diffuser FR First curved surface SR Second curved surface IP Inflection point PC Flow path center

Claims (4)

In a compressor for a jet engine that compresses air taken in an annular core flow path formed in a cylindrical engine case in a jet engine,
A cylindrical compressor case constituting a part of the engine case;
It is provided in the compressor case so as to be rotatable along the axial direction, and is provided integrally at equal intervals on the outer peripheral surface of the disk, which is rotatable around the axis, and located in the core flow path. A multi-stage compressor rotor with a plurality of rotor blades;
A plurality of compressor blades provided alternately with a plurality of stages of compressor rotors along the axial direction in the compressor case, and provided with a plurality of stationary blades disposed at equal intervals in the circumferential direction and positioned in the core flow path A compressor stator of a stage,
Each stator blade in the compressor stator at the final stage is each stator blade of the compressor stator at the other stage so that each stator blade of the compressor stator at the final stage has a function of supporting the compressor case. Compared to the above, the length of the cord is longer, and a part of the core flow path is formed between the compressor stator at the final stage and the combustor in the jet engine, and the compressed air is decelerated to reduce the combustor. An annular diffuser is formed
A portion between the leading edge side and the trailing edge side of the stationary blade of the compressor stator at the final stage on the radially inner wall surface of the core channel transitions from a concave first curved surface to a convex second curved surface. The inflection point at which the first curved surface transitions to the second curved surface is a point separated by 70 to 85% code length from the leading edge of the stationary blade of the compressor stator at the final stage. Compressor for jet engine. The leading edge of the stationary blade of the compressor stator at the final stage is a tip on the hub side when the flow path center of the diffuser gradually moves away from the axis of the compressor case toward the downstream side. When the center of the flow path of the diffuser is gradually approaching the axial center of the compressor case toward the downstream side, the hub side is more than the tip side. The jet engine compressor according to claim 1, wherein the compressor is swept so as to be positioned upstream . The camber angle of the stationary blade of the compressor stator at the final stage is such that the flow path center of the diffuser gradually moves away from the axis of the compressor case toward the downstream side. Gradually decreases from the tip side to the tip side, and when the flow path center of the diffuser gradually approaches the axial center of the compressor case toward the downstream side, it gradually increases from the center span side to the hub side. 3. The jet engine compressor according to claim 1, wherein the compressor is small . Jet engine, characterized by comprising a compressor for a jet engine according to any one of claims 1 to 3.

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