JP2008501091A - Gas turbine compression system and compressor structure - Google Patents

Abstract

The invention is arranged between a gas channel, a low pressure compressor part (8) and a high pressure compressor part (9) for compressing the gas in the channel, and a low pressure compressor part (8) and a high pressure compressor part (9). It relates to a gas turbine compression system (1) having a compressor structure (14). The compressor structure (14) is configured to guide the gas flow in the gas channel and has a plurality of radial struts (15, 16, 21, 24, 25) for transmitting loads, At least one of the struts (15, 16, 21, 24, 25) is hollow to accommodate service parts. The compressor structure (14) is arranged immediately downstream of the last rotor (10) in the low-pressure compressor section (8) and has a plurality of the struts (15, 16, 21, 24, 25 having a camber shape). ) To greatly change the direction of gas vortex flow from the rotor (10).
[Selection] Figure 1

Description

  The present invention has a gas channel, a low-pressure compressor unit and a high-pressure compressor unit for compressing gas in the channel, and a compressor structure unit disposed between the low-pressure compressor unit and the high-pressure compressor unit, A gas configured to guide a gas flow in the gas channel and having a plurality of radial struts for transmitting a load, wherein at least one of the struts is hollow to accommodate a service component The present invention relates to a turbine compression system. The invention also relates to a compressor structure.

  The gas turbine compression system forms part of a gas turbine engine. The gas turbine compression system is in a gas turbine engine driven by the turbine system via the engine shaft. Gas turbine engines are particularly for aircraft jet engines. Jet engines include various types of engines that accept air at a relatively low speed, heat it by combustion, and release it at a much higher speed. The expression of the jet engine includes, for example, a turbo jet engine and a turbo fan engine. The present invention is described below for a turbofan engine, but of course can be used for other engine types.

  The structural strength of a gas turbine engine depends on a limited number of engine structural parts, also known as cases. Thus, the structure represents the skeleton of the engine. The structure is subject to high loads during engine operation. The structure typically has a bearing housing for the engine shaft, a gas flow channel in the form of an annular duct, and a radial strut that forms a link between the inner and outer portions of the engine. The compressor structure of the present invention forms such a structure.

  Struts are often hollow to accommodate service components such as means for oil and / or air intake and exhaust, and to accommodate instruments such as electrical and metal cables for the transmission of information regarding measured pressure and / or temperature, etc. It is. In order to minimize the effect on the gas flow, the struts are usually airfoil shaped with a symmetrical cross section. Generally, service requirements determine the number of struts required.

  The primary objective of the present invention is to reduce the number of parts in a gas turbine compression system.

  The purpose of this is to arrange the compressor structure part immediately downstream of the last rotor in the low-pressure compressor part and to change the direction of the gas vortex from the rotor by a plurality of the struts having a camber shape. Achieved. “Changing the direction significantly” means changing the direction of the gas flow by at least 20 °. Furthermore, the gas vortex is changed in a direction in which the dominant component becomes the axial direction. The compressor structure may be configured to change the direction of gas flow in a direction substantially parallel to the rotational axis of the engine.

  In conventional gas turbine compression systems, there is a last stator row between the last rotor and strut in the low pressure compressor section. This last stator row has a plurality of aerodynamic vanes (often about 150 vanes) configured to change the direction of the gas vortex flow from the last rotor in the low pressure compressor section to a substantially axial direction. According to the present invention, the last stator row can be removed. In other words, according to the present invention, the function of the last stator row and the function of the conventional compressor structure with struts is replaced by the compressor structure of the present invention.

  Furthermore, conventional compressor structures with struts having a symmetrical airfoil shape have a very limited aerodynamic function. This represents a “dead load” from an aerodynamic perspective. Conventional compressor structures are substantially detrimental to pressure loss. In addition, the length of the gas flow channel through the compressor structure is generally determined by aerodynamic constraints of gentle axial-to-radial flow diversion or to the size of the bearing housing to avoid boundary layer peeling. The Both of these two constraints make the gas flow channel relatively long without taking advantage of the part length available from an aerodynamic perspective, which affects engine length. Therefore, according to the present invention, by using a substantially aerodynamically shaped strut, the aerodynamic functionality of the compressor structure is not expected to have an adverse effect on the required overall engine length.

  In other words, according to the present invention, the available axial length of the compressor structure determined by the overall engine layout and aerodynamic constraints is used to strut the aerodynamic functionality of the upstream stator train. It can be integrated into the compressor structure.

  While reducing the number of engine parts, it is possible to reduce, or at least not increase, the flow distortion for the downstream high pressure compressor rotor and the upstream low pressure compressor rotor.

  To achieve a significant change in gas flow orientation, the camber centerline direction at at least one leading edge of the camberast rat is at least 20 ° relative to the camber centerline direction at the trailing edge of the camberat rat. Inclined.

  According to a preferred embodiment of the present invention, the at least one thickness to chord ratio of the canburst rat is about 0.10 (± 0.05). This ratio can be optimized with respect to the number of aerodynamic vanes and struts. Thereby, the state which made the cross-sectional area of each strut larger arises. Increasing the cross-sectional area results in an increase in structural strength and service capacity per strut. This can be chosen to reduce the total number of load bearing struts and / or increase the overall service capacity of the compressor structure.

  According to a further development, the compressor structure comprises a plurality of aerodynamic vanes having a cross-sectional area substantially smaller than the struts. These smaller aerodynamic vanes are positioned to help the struts change flow direction. The aerodynamic vanes may alternatively be arranged to produce a favorable pressure distribution around the struts.

  According to a further embodiment, the struts are distributed asymmetrically in the circumferential direction of the compressor structure. This creates a state in which the struts are distributed so as to obtain the optimum structural strength.

  The present invention will now be described with reference to the embodiments shown in the accompanying drawings.

  The present invention will be described below with reference to FIG. 1 for an aircraft engine 1 with a high bypass ratio. The engine 1 has an outer housing 2, an inner housing 3 and an intermediate housing 4, which is concentric with the first two housings and has a gap between them as an inner primary gas channel 5 for compressing propulsion gas. And the secondary channel 6 through which the engine bypass circulates. Therefore, each of the gas channels 5 and 6 is annular in a cross section perpendicular to the axial direction 18 of the engine 1. A fan 7 is arranged at the engine intake upstream of the inner and outer gas channels 5, 6.

  The engine 1 includes a low-pressure compressor unit 8 and a high-pressure compressor unit 9 for compressing the gas in the primary gas channel 5. In order to burn the compressed gas from the primary gas channel 5, a combustion chamber 17 is disposed on the downstream side of the high-pressure compressor unit 9. The aircraft engine 1 further comprises a compressor section (not shown) for expanding the propellant gas, which is arranged downstream of the combustion chamber as is known in the art.

  Each compressor part 8, 9 of the compressor has a plurality of rotors 10, 11 and stators 12, 13 between two adjacent rotors. The stators 12 and 13 have a plurality of aerodynamic vanes for changing the gas vortex from the upstream rotor into a substantially axial direction.

  The housings 2, 3, 4 are supported by structural parts 14, 15 that connect the housings by radial arms. These arms are commonly known as struts. The struts must provide this support and be sufficiently resistant so that they will not break or buckle when the fan blades loosen and collide with them. Furthermore, the strut is configured to transmit a load in the engine. In addition, the struts are used to house service components such as means for intake and discharge of oil and / or air, and to drive appliances such as electrical and metal cables for transmission of information on measured pressure and / or temperature and for start engines. It is hollow to accommodate the shaft. The struts can also be used for coolant guidance.

  The compressor structure 14 that connects the intermediate housing 4 and the inner housing 3 is conventionally called an intermediate case (IMC) or an intermediate compressor case (ICC). The compressor structure portion 14 is configured to guide the gas flow from the low pressure compressor portion 8 inward in the radial direction toward the high pressure compressor portion inlet. The compressor structure 14 connecting the intermediate housing 4 and the inner housing 3 has a plurality of radial struts spaced from each other in the circumferential direction of the compressor structure 14 as shown in FIGS. 15, 16, 21, 24, 25. These struts 15, 16 are structural parts configured to transmit both axial and radial loads and are hollow to accommodate service parts.

  The compressor structure 14 is configured to change the direction of the gas vortex flow from the rotor 10 to substantially the axial direction. Therefore, the compressor structure portion 14 is arranged immediately downstream of the last rotor 10 in the low-pressure compressor portion 8. Further, the compressor structure portion 14 is disposed immediately upstream of the first rotor 11 in the high-pressure compressor portion 9. The gas vortex from the rotor 10 usually flows at an angle of 40-60 ° with respect to the axial direction 18 of the engine. The struts 15 and 16 are disposed immediately downstream of the last rotor 10 in the low-pressure compressor unit 8. In this case, the change in the direction of the gas flow is a coupling direction of axial direction-tangential direction and axial direction-radial direction.

  The magnitude of the change in the direction of the gas flow in the compressor structure 14 depends on several parameters. In order to change the direction of the gas flow with a size of 40-60 °, the struts 15, 16, 21, 24, 25 have a camber airfoil shape, as shown in FIGS. In other words, the strut is configured with a sufficient curvature to greatly change the direction of gas flow. Thus, the struts 15, 16, 21, 24, 25 are not only structural but also aerodynamic. More specifically, the direction of the camber centerline M at the leading edge 101 of the camberast rat 16 is at an angle corresponding to the desired turning angle with respect to the direction of the camber centerline M at the trailing edge 102 of the camburst rat. Inclined. Therefore, the direction of the camber centerline M at the leading edge 101 of the camberast rat 16 is at least 20 °, suitably at least 30 °, in particular at least 40 relative to the direction of the camber centerline M at the trailing edge 102 of the camburst rat. Incline, preferably at least 50 °.

  In order to change the direction of gas flow to a size of 40-60 °, the struts are further configured with longer chords compared to conventional struts. The chord is defined in FIG. 5 as the distance between the leading edge 101 and trailing edge 102 of the vane 15 along the chord line C. The chord line C is defined as a straight line connecting the leading edge 101 and the trailing edge 102. More specifically, the chords of the canburst rats 15, 16, 21, 24, 25 are at least 6 times the thickness of the canburst rats, suitably at least 7 times, preferably at least 8 times, in a preferred embodiment. According to it is about 9 times. On the other hand, the strut thickness may be approximately the same as a conventional strut.

  The strut thickness is defined as the maximum distance between two opposing strut surfaces 103, 104 in a direction perpendicular to the camber centerline M. The camber centerline M is defined as the trajectory of the midpoint between the upper and lower surfaces of the strut when measured perpendicular to the camber centerline itself. Camber A is defined as the maximum distance between the camber center line M and the chord line C, measured perpendicular to the chord line. In accordance with the present invention, the strut chord is considerably longer than the conventional strut chord.

  Further, maximum thickness to chord ratio is another measure of strut gas flow diverting capability. The maximum thickness is preferably less than 20%, in particular less than 15%, more specifically about 10% of the chord according to the example shown in the drawing.

  The struts 15, 16, 21, 24, 25 are further distributed asymmetrically in the circumferential direction of the annular compressor structure 14, as shown in FIG. The first of these struts 15 is located in the highest possible vertical position in the gas channel of the compressor structure. The first strut 15 is somewhat thicker than the other struts 16, 21, 24, 25 to receive the radial drive shaft for the start engine. The further struts 16, 21 and 24, 25 are each distributed symmetrically with respect to a plane 23 which coincides with the first strut 15 and which is parallel to the axial direction 18 of the gas turbine compression system. More specifically, the two struts 16, 21 and 24, 25 are respectively arranged on each side of the symmetry plane 23.

  A plurality of so-called aerodynamic or splitter vanes 19, 20 are arranged between the struts 15, 16, 21, 24, 25. Thus, the aerodynamic vanes 19, 20 are arranged in the compressor structure 14 and form a single circular cascade with the struts. Aerodynamic vanes are significantly smaller and lighter than struts and are non-bearing from a structural point of view. The number of struts is much less than aerodynamic vanes.

  The aerodynamic vanes 19, 20 are arranged to help the struts change the direction of gas flow from the rotor 10 in a substantially axial direction.

  Splitter vanes 19, 20 are positioned and staggered to reduce the risk of pressing applied to the upstream rotor train from the non-rotating flow of the struts. The strut profile is also optimized to mitigate the upstream impact of the non-rotating flow of the strut on the upstream rotor by selection of the appropriate leading edge radius and wedge angle.

  In addition, splitter vanes 19, 20 that do not have a structural support function reduce the strength of secondary flows such as U-shaped and passage vortices by controlling the flow rate and pressure gradient in the low aspect ratio strut passages. It can be extended in an arcuate direction in the chord direction, inclined, or even bent in the cross chord direction.

  Also, the need for splitter vanes to reduce the number of struts allows the introduction of the asymmetric strut arrangement in the circumferential direction within the IMC / ICC duct as shown in FIG. It becomes possible. In fact, the struts are arranged to receive structural loads, in particular engine mount loads, in an optimal manner from a structural point of view. Thus, the splitter vanes 19, 20 are distributed so as to uniformly remove the flow vortices despite the asymmetric distribution of struts.

  In the above description, the rotational axis of the engine and the axial direction of the engine / gas turbine compression system / compressor structure point to the same shaft 18.

  The present invention is in no way limited to the embodiments described above, and many alternatives and modifications are possible without departing from the scope of the appended claims.

  For example, the smaller aerodynamic vane (splitter vane) arrangement of FIG. 4 is only one illustration of a possible configuration, whereby the splitter vane is positioned axially, radially, and tangentially with respect to the struts. Is not perfect as to. As an example, an aerodynamic vane may be placed at the trailing edge of a particular strut to form a type of flap to increase the strut's ability to redirect gas flow.

  Instead of being solid, the aerodynamic vane may have a hollow cross section, i.e. have at least substantial voids / cavities, but not necessarily through holes.

  According to an alternative embodiment, the compressor structure may be configured to change the direction of gas flow in a direction different from the axial direction. The compressor structure may be configured, for example, to change the gas flow from a + 50 ° inflow direction to a −10 ° outflow direction. The maximum gas flow diverting capability of the compressor structure will be about 60-70 °.

It is a schematic side view of the engine cut along a plane parallel to the rotation axis of the engine. It is an enlarged view of the compressor structure part between the low pressure compressor part and high pressure compressor part of FIG. It is sectional drawing along AA of FIG. It is an expanded sectional view along BB of FIG. FIG. 5 is an enlarged cross-sectional view of one of the struts of FIG. 4.

Claims (25)

A gas channel (5), a low-pressure compressor section (8) and a high-pressure compressor section (9) for compressing the gas in the channel, and a compressor arranged between the low-pressure compressor section (8) and the high-pressure compressor section (9) And the compressor structure (14) is configured to guide the gas flow in the gas channel and also includes a plurality of radial struts (15, 16, 21) for transmitting loads. 24, 25), wherein at least one of the struts is hollow to accommodate a service component, the gas turbine compression system (1) comprising:
The compressor structure (14) is arranged immediately downstream of the last rotor (10) in the low-pressure compressor section (8) and has a plurality of the struts (15, 16, 21, 24, 25) having a camber shape. The gas turbine compression system is configured to greatly change the direction of the gas vortex from the rotor (10).   The direction of the camber center line (M) at at least one leading edge (101) of the camburst rat (15, 16, 21, 24, 25) is the camber center line (M The gas turbine compression system according to claim 1, wherein the gas turbine compression system is inclined at least 20 ° with respect to the direction of.   The direction of the camber center line (M) at at least one leading edge (101) of the camburst rat (15, 16, 21, 24, 25) is the camber center line (M The gas turbine compression system according to claim 1, wherein the gas turbine compression system is inclined at least 30 ° with respect to the direction of.   The direction of the camber center line (M) at at least one leading edge (101) of the camburst rat (15, 16, 21, 24, 25) is the camber center line (M 2. The gas turbine compression system according to claim 1, wherein the gas turbine compression system is inclined at least 40 ° relative to the direction of   Gas turbine according to any of the preceding claims, characterized in that at least one thickness to chord ratio of the canburst rat (15, 16, 21, 24, 25) is about 0.10. Compression system.   The struts (15, 16, 21, 24, 25) are characterized in that their respective leading edges (101) are arranged in substantially the same position in a direction parallel to the rotation axis (18) of the engine. A gas turbine compression system according to any one of claims 1 to 5.   The compressor structure (14) comprises a plurality of aerodynamic vanes (19, 20) having a substantially smaller cross-sectional area than the struts (15, 16, 21, 24, 25). A gas turbine compression system according to any one of the above.   The at least one aerodynamic vane (19, 20) is located between two adjacent struts (16, 21) in the circumferential direction of the compressor structure (14). Gas turbine compression system.   The leading edge of each of the first set of aerodynamic vanes (19, 20) is substantially the same position as the leading edge of the struts (15, 16, 21, 24, 25) in a direction parallel to the rotational axis (18) of the engine. The gas turbine compression system according to claim 7 or 8, wherein   The gas turbine compression system according to any of claims 7 to 9, characterized in that at least one of the aerodynamic vanes (19, 20) is substantially solid in cross section.   Gas turbine compression according to any one of the preceding claims, characterized in that the struts (15, 16, 21, 24, 25) are distributed asymmetrically in the circumferential direction of the compressor structure. system.   12. The strut (15, 16, 21, 24, 25) is distributed symmetrically with respect to a plane parallel to the axial direction of the gas turbine compression system. Gas turbine compression system.   13. A gas turbine compression system according to any of the preceding claims, characterized in that the compressor structure (14) is arranged immediately upstream of the first rotor (11) in the high-pressure compressor section. Having a gas channel (5) and a plurality of radial struts (15, 16, 21, 24, 25) for transmitting a load, at least one of the struts being hollow to accommodate service parts A compressor structure (14) configured to guide a gas flow during operation of a gas turbine compression system (1),
When the compressor structure is disposed in a gas turbine compression system, the plurality of struts (15, 16, 21, 24, 25) having a camber shape direct the direction of the gas vortex from the upstream rotor (10). Compressor structure, characterized in that it is configured to change significantly.   The direction of the camber center line (M) at at least one leading edge (101) of the camburst rat (15, 16, 21, 24, 25) is the camber center line (M The compressor structure according to claim 14, wherein the compressor structure is inclined at least 20 ° with respect to the direction of.   The direction of the camber center line (M) at at least one leading edge (101) of the camburst rat (15, 16, 21, 24, 25) is the camber center line (M The compressor structure according to claim 14, wherein the compressor structure is inclined at least 30 ° with respect to the direction of.   The direction of the camber center line (M) at at least one leading edge (101) of the camburst rat (15, 16, 21, 24, 25) is the camber center line (M The compressor structure according to claim 14, wherein the compressor structure is inclined at least 40 ° with respect to the direction of.   18. Compressor structure according to any of claims 14 to 17, characterized in that at least one thickness to chord ratio of the canburst rat (15, 16, 21, 24, 25) is about 0.10. Department.   The struts (15, 16, 21, 24, 25) are arranged such that their respective leading edges (101) are substantially in the same position in a direction parallel to the central axis (18) of the compressor structure. The compressor structure according to any one of claims 14 to 18.   The compressor structure (14) comprises a plurality of aerodynamic vanes (19, 20) having a substantially smaller cross-sectional area than the struts (15, 16, 21, 24, 25). The compressor structure part in any one of.   21. The at least one aerodynamic vane (19, 20) is located between two adjacent struts (16, 21) in the circumferential direction of the compressor structure (14). Compressor structure part.   The leading edge of each of the first set of aerodynamic vanes (19, 20) is approximately the leading edge of the struts (15, 16, 21, 24, 25) in a direction parallel to the central axis (18) of the compressor structure. The compressor structure according to claim 20 or 21, wherein the compressor structure is disposed at the same position.   23. Compressor structure according to any of claims 20 to 22, characterized in that at least one of the aerodynamic vanes (19, 20) is substantially solid in cross section.   24. A compressor structure according to any one of claims 14 to 23, characterized in that the struts (15, 16, 21, 24, 25) are distributed asymmetrically in the circumferential direction of the compressor structure. .   25. A strut according to any one of claims 14 to 24, characterized in that the struts (15, 16, 21, 24, 25) are distributed symmetrically with respect to a plane parallel to the central axis of the compressor structure. Compressor structure.

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