A gas turbine compressor vortex finder mechanism

Abstract

This invention discloses a vortex reducer mechanism for a gas turbine compressor, belonging to the field of aero-engines. It includes an axial bleed section, a drum, and opposing front-stage and rear-stage compressor disks. The two ends of the drum are fixed to the inner sides of the front-stage and rear-stage compressor disks, respectively. The front-stage, rear-stage, and drum disks, along with the drum, form a disk cavity that communicates with the axial bleed section. Multiple drum holes are formed on the outer circumferential surface of the drum, each communicating with the disk cavity. An annular partition is provided between the front-stage and rear-stage compressor disks, and multiple drainage holes corresponding to the drum holes are formed on the outer circumferential surface of the partition. Multiple sets of flow guiding components are provided between the front-stage and rear-stage compressor disks, each corresponding to one of the drainage holes. This invention effectively reduces the significant pressure loss generated by the cold air.

Description

Technical Field

[0001] This invention relates to the field of aero engines, and more particularly to a gas turbine compressor vortex reducer mechanism. Background Technology

[0002] With the development of aero-engine and gas turbine design technologies and the continuous advancement of requirements, the temperature and thermal stress of hot-end components are becoming increasingly higher, placing higher demands on the design technology of the internal air system of aero-engines. Currently, internal cooling air is mainly drawn from the compressor platform, flows through the drum orifice, and then radially enters the common disk cavity. Through the inter-shaft channel, it is transported to hot-end components and other places to complete cooling and sealing functions before merging into the main combustion stream or being discharged into the atmosphere.

[0003] Currently, most aero-engines use a method of opening holes in the compressor drum between the compressor discs to radially introduce cooling airflow and then axially guide it out to cool hot-end components. During the radial inward flow of the airflow, due to the Coriolis effect of the compressor rotating disc cavity, a large-scale free vortex structure is generated after the airflow enters the drum hole, which causes a significant pressure loss of the cooling air. Summary of the Invention

[0004] In order to improve the problem of large-scale free vortex structure generated after the airflow enters the drum orifice, which causes a large pressure loss, this application provides a gas turbine compressor vortex reducer mechanism.

[0005] The gas turbine compressor vortex reducer mechanism provided in this application adopts the following technical solution:

[0006] A gas turbine compressor vortex reducer mechanism includes an axial bleed section, a drum, and a front-stage compressor disk and a rear-stage compressor disk arranged opposite to each other. The two ends of the drum are respectively fixed to the inner sides of the front-stage compressor disk and the rear-stage compressor disk. The front-stage compressor disk, the rear-stage compressor disk, and the drum form a disk cavity, which is connected to the axial bleed section. The outer circumferential surface of the drum has multiple drum holes, each communicating with the disk cavity. An annular partition is provided between the front-stage compressor disk and the rear-stage compressor disk. The outer circumferential surface of the annular partition has multiple drainage holes corresponding to the multiple drum holes. Multiple sets of flow guiding components are provided between the front-stage compressor disk and the rear-stage compressor disk. These multiple sets of flow guiding components are located between the central axis of the annular partition and the front-stage compressor disk, and each set of flow guiding components corresponds to one of the multiple drainage holes.

[0007] By adopting the above technical solution, the cooling gas enters the disk cavity through the drum hole, and then the gas in the disk wall passes through the drainage hole. The gas passing through the drainage hole is guided by the flow guiding component, so that the cold gas passes through the axial air intake section, thereby reducing the formation of free vortex structure in the cooling gas in the disk cavity, and thus reducing the large pressure loss of the cooling gas.

[0008] Preferably, each group of the flow guiding components includes multiple flow guide plates one and multiple flow guide plates two, which are arranged opposite to each other. The ends of the flow guide plates one and two that are far apart from each other are respectively arranged on the side opposite to the front stage compressor disk and the rear stage compressor disk. The flow guide plates one and two are respectively inclined toward the central axis of the front stage compressor disk.

[0009] By adopting the above technical solution, when the gas in the disk cavity passes through the drainage hole, the cooling gas passes between multiple guide plates one and multiple guide plates two, so that multiple guide plates one and multiple guide plates two guide the cooling gas, thereby allowing the cooling gas in the disk cavity to pass through the axial air intake section, reducing the formation of free vortex structure of cooling gas in the disk cavity.

[0010] Preferably, both the first and second guide vanes are arc-shaped plates, with the arc surfaces of the first and second guide vanes protruding away from the central axis of the front-stage compressor disk, and the areas of the first and second guide vanes gradually decreasing towards the central axis of the front-stage compressor disk.

[0011] By adopting the above technical solution, multiple guide plates 1 and multiple guide plates 2 are set as arc-shaped plates, and the arc-shaped plates bulge in a direction away from the central axis of the front stage compressor disk, thereby reducing the resistance of guide plates 1 and 2 to the cooling gas blown out of the duct hole, and facilitating the cooling gas to pass through the axial duct section.

[0012] Preferably, the sides of the multiple guide vanes one away from the central axis of the front stage compressor disk and the sides of the multiple guide vanes two away from the central axis of the front stage compressor disk are all formed in moving slots. A push rod is inserted into the multiple moving slots. A sliding slot is formed on the opposite inner side of each moving slot. A sliding rod is slidably arranged in the sliding slot. One end of the sliding rod near the central axis of the push rod is fixedly connected to the outer peripheral surface of the push rod. The sides of the front stage compressor disk and the rear stage compressor disk opposite to each other are respectively provided with a driving mechanism for driving the push rod to move radially along the front stage compressor disk.

[0013] By adopting the above technical solution, the push rod is driven by the drive mechanism to move radially along the front stage compressor disk, so that the push rod drives the sliding rod to move radially along the front stage compressor disk, thereby causing multiple guide vanes one and multiple guide vanes two to swing, adjusting the tilt angle of guide vanes one and guide vanes two, and thus adjusting the flow rate of cooling gas in the disk cavity.

[0014] Preferably, the drive mechanism includes multiple connecting plates fixed to the opposite sides of the front-stage compressor disc and the rear-stage compressor disc, each connecting plate corresponding to a plurality of push rods. A sleeve is provided through the connecting plate, the sleeve is rotatably connected to the connecting plate, the push rod passes through the sleeve, and the push rod is threadedly connected to the sleeve. The opposite sides of the front-stage compressor disc and the rear-stage compressor disc are respectively provided with drive components for driving the multiple sleeves to rotate.

[0015] By adopting the above technical solution, multiple sleeves are simultaneously driven to rotate by the drive component, causing the sleeve drive push rod to move along the axial direction of the sleeve, thereby causing the push rod to drive the sliding rod to move along the axial direction of the sleeve, adjusting the tilt angle of multiple guide plates one and multiple guide plates two.

[0016] Preferably, the drive assembly includes a gear 1 disposed on the outer peripheral surface of the sleeve, and a gear ring is rotatably disposed on the opposite side of the front stage compressor disk and the rear stage compressor disk, the gear ring meshing with a plurality of the gear 1s, and a drive member for driving the two gear rings to rotate is respectively disposed on the opposite side of the front stage compressor disk and the rear stage compressor disk.

[0017] By adopting the above technical solution, the gear ring is driven to rotate by the driving component, and the gear ring simultaneously drives multiple gears to rotate, thereby causing multiple sleeves to rotate simultaneously.

[0018] Preferably, the drive unit includes two motors, which are respectively fixed on the sides of the front stage compressor disk and the rear stage compressor disk opposite to each other. A second gear is provided on the output shaft of the motor, and the second gear meshes with the gear ring.

[0019] By adopting the above technical solution, the motor is started, the output shaft of the motor drives gear two to rotate, gear two drives the gear ring to rotate, the gear ring drives multiple gear one to rotate, thereby causing multiple gear one to drive multiple sleeves to rotate simultaneously.

[0020] Preferably, an annular groove 1 is provided on the opposite sides of the front stage compressor disk and the rear stage compressor disk, and a circular ring 1 is fixed on the opposite sides of the two gear rings. The circular ring 1 is inserted into the annular groove 1. An annular groove 2 is provided on the inner circumferential surface of the annular groove 1. A circular ring 2 is fixed on the outer circumferential surface of the circular ring 1. The circular ring 2 is disposed in the annular groove 2.

[0021] By adopting the above technical solution, when the gear ring rotates, the gear ring drives the first ring to rotate, and the first ring drives the second ring to rotate in the second annular groove, thereby making the rotation of the gear ring more stable.

[0022] Preferably, a guide ring is fixed to the side of the front-stage compressor disk near the rear-stage compressor disk, the central axis of the guide ring is collinear with the central axis of the front-stage compressor disk, and the outer peripheral surface of the guide ring is an arc-shaped surface.

[0023] By adopting the above technical solution, a guide ring is set on the side of the front stage compressor disk near the rear stage compressor disk. When the gas in the disk cavity passes through the outer circumference of the guide ring between multiple guide plates one and multiple guide plates two, the guide ring guides the cooling gas in the disk cavity, making it easier for the cooling gas to pass through the axial air intake section.

[0024] Preferably, a plurality of guide vanes are fixed on the outer circumferential surface of the guide ring, and the plurality of guide vanes are arranged at equal intervals along the circumference of the guide ring.

[0025] By adopting the above technical solution, when the gas in the disk cavity blows into the guide ring, the cooling gas is diverted by the guide vanes, so that the cooling gas is evenly distributed circumferentially when passing through the axial air intake section.

[0026] In summary, this application includes at least one of the following beneficial technical effects:

[0027] 1. Cooling gas enters the disk cavity through the drum hole, and then the gas in the disk wall passes through the drainage hole. The gas passing through the drainage hole is guided by the flow guiding component, so that the cold gas passes through the axial air intake section, thereby reducing the formation of free vortex structure in the cooling gas in the disk cavity, and thus reducing the large pressure loss of the cooling gas.

[0028] 2. When the gas in the disk cavity passes through the drainage hole, the cooling gas passes between multiple guide plates one and multiple guide plates two, so that multiple guide plates one and multiple guide plates two guide the cooling gas, thereby allowing the cooling gas in the disk cavity to pass through the axial air intake section, reducing the formation of free vortex structure of cooling gas in the disk cavity;

[0029] 3. The push rod is driven to move radially along the front stage compressor disk by the drive mechanism, so that the push rod drives the sliding rod to move radially along the front stage compressor disk, thereby causing multiple guide vanes one and multiple guide vanes two to swing, adjusting the tilt angle of guide vanes one and two, and thus adjusting the flow rate of cooling gas in the disk cavity. Attached Figure Description

[0030] Figure 1 This is a schematic diagram of the overall structure of the gas turbine compressor vortex reducer mechanism in an embodiment of this application.

[0031] Figure 2 These are cross-sectional views of the front-stage compressor disk and the rear-stage compressor disk in the embodiments of this application.

[0032] Figure 3 This is a cross-sectional view of a guide vane in one of the embodiments of this application.

[0033] Figure 4 yes Figure 3 Enlarged diagram of point A in the middle.

[0034] Figure 5 yes Figure 2 Enlarged diagram of point B in the middle.

[0035] Figure 6 This is a cross-sectional view of the rear stage compressor disc in an embodiment of this application.

[0036] Figure 7 yes Figure 6 Enlarged diagram of point C in the middle.

[0037] Reference numerals: 1. Axial air intake section; 2. Front stage compressor disc; 3. Rear stage compressor disc; 4. Drum; 41. Drum orifice; 42. Disc cavity; 5. Annular baffle; 51. Drain hole; 6. Flow guide assembly; 61. Flow guide plate one; 62. Flow guide plate two; 63. Moving groove; 64. Sliding groove; 65. Push rod; 66. Sliding rod; 7. Connecting plate; 71. Sleeve; 72. Gear one; 73. Gear ring; 74. Annular groove one; 75. Circular ring one; 76. Annular groove two; 77. Circular ring two; 78. Motor; 79. Gear two; 8. Flow guide ring; 81. Flow guide vane. Detailed Implementation

[0038] The following is in conjunction with the appendix Figure 1-7 This application will be described in further detail.

[0039] This application discloses a gas turbine compressor vortex reducer mechanism.

[0040] Reference Figure 1 and Figure 2 A gas turbine compressor vortex reducer mechanism includes an axial induced draft section 1 and a front-stage compressor disk 2 and a rear-stage compressor disk 3 arranged opposite to each other. A drum 4 is disposed between the front-stage compressor disk 2 and the rear-stage compressor disk 3. The two end faces of the drum 4 are fixedly connected to the opposite sides of the front-stage compressor disk 2 and the rear-stage compressor disk 3. The outer peripheral surface of the drum 4 is coplanar with the outer peripheral surfaces of the front-stage compressor disk 2 and the rear-stage compressor disk 3. The front-stage compressor disk 2, the rear-stage compressor disk 3, and the drum 4 together form a disk cavity 42. The outer peripheral surface of the drum 4 has a plurality of drum holes 41, which are equally spaced along the circumference of the drum 4.

[0041] Reference Figure 1and Figure 2 An annular partition 5 is fixed between the front-stage compressor disk 2 and the rear-stage compressor disk 3. The central axis of the annular partition 5 is collinear with the central axis of the front-stage compressor disk 2. Multiple drainage holes 51 are provided on the outer circumferential surface of the annular partition 5. The multiple drainage holes 51 correspond to multiple drum holes 41 and are evenly spaced along the circumference of the annular partition 5.

[0042] Reference Figure 2 and Figure 3 Multiple sets of flow guiding components 6 are arranged between the front-stage compressor disk 2 and the rear-stage compressor disk 3. These components are located between the annular partition 5 and the central axis of the front-stage compressor disk 2. Each set of flow guiding components 6 corresponds to multiple flow inlets 51. The multiple flow guiding components 6 include multiple guide plates 61 and multiple guide plates 62. Both guide plates 61 and 62 are arc-shaped plates, with their arc surfaces bulging away from the central axis of the front-stage compressor disk 2. The guide plates 61 and 62 are arranged opposite each other, with the guide plates 61 arranged at equal intervals radially along the front-stage compressor disk 2 and the guide plates 62 arranged at equal intervals radially along the rear-stage compressor disk 3.

[0043] Reference Figure 2 Multiple guide vanes 61 are hinged at one end away from multiple guide vanes 62 to the side of the front stage compressor disk 2 near the rear stage compressor disk 3. Multiple guide vanes 62 are hinged at one end away from multiple guide vanes 61 to the side of the rear stage compressor disk 3 near the front stage compressor disk 2. The bottom ends of multiple guide vanes 61 and multiple guide vanes 62 are inclined toward the central axis of the front stage compressor disk 2. The area of ​​multiple guide vanes 61 and multiple guide vanes 62 gradually decreases toward the central axis of the front stage compressor disk 2.

[0044] Cooling gas is introduced into the disk cavity 42 through the drum hole 41. Then, the cold gas in the disk cavity 42 passes through the drainage hole 51 and is guided and diverted by multiple guide plates 61 and 62, so that the cold gas passes through the axial air intake section 1, reducing the formation of a free vortex structure in the disk cavity 42, thereby reducing the large pressure loss of the cold gas.

[0045] Reference Figure 2 , Figure 3 and Figure 4Multiple guide vanes 61 and multiple guide vanes 62 are provided with moving grooves 63 on their sides away from the central axis of the front stage compressor disk 2. Push rods 65 are inserted into the multiple moving grooves 63. A sliding groove 64 is provided on one of the opposite inner sides of the moving grooves 63. A sliding rod 66 is slidably arranged in the sliding groove 64. The end face of the sliding rod 66 away from the push rod 65 is in contact with the inner side of the sliding groove 64 away from the push rod 65. The end of the sliding rod 66 near the central axis of the push rod 65 is fixedly connected to the outer peripheral surface of the push rod 65.

[0046] Reference Figure 2 , Figure 3 and Figure 5 Multiple connecting plates 7 are fixed to the opposite sides of both the front-stage compressor disc 2 and the rear-stage compressor disc 3. These connecting plates 7 are located on the side of the multiple guide vanes 61 away from the annular partition 5, and correspond to multiple push rods 65. A sleeve 71 passes through the top surface of each connecting plate 7, and the sleeve 71 is rotatably connected to the connecting plate 7. One end of the push rod 65 near the central axis of the front-stage compressor disc 2 passes through the sleeve 71, and the push rod 65 is threadedly connected to the sleeve 71. A gear 72 is fixedly fitted onto the outer circumference of the sleeve 71. Gear rings 73 are rotatably mounted on the opposite sides of both the front-stage compressor disc 2 and the rear-stage compressor disc 3, and these gear rings 73 mesh with the multiple gears 72. Both the front-stage compressor disk 2 and the rear-stage compressor disk 3 have annular grooves 74 on their opposite sides. Two gear rings 73 have circular rings 75 fixed to their respective distant sides. The central axis of the circular rings 75 is collinear with the central axis of the gear rings 73, and the circular rings 75 are inserted into the annular grooves 74. Annular grooves 76 are formed on the inner circumferential surface of the annular grooves 74 away from the central axis of the front-stage compressor disk 2. A circular ring 77 is fixed to the outer circumferential surface of the circular rings 75, and the circular ring 77 is disposed within the annular groove 76.

[0047] Reference Figure 6 and Figure 7 Motors 78 are fixed on opposite sides of the front stage compressor disk 2 and the rear stage compressor disk 3. A gear 79 is fixed on the output shaft of the motor 78, and the gear 79 meshes with the gear ring 73.

[0048] Reference Figure 2 A guide ring 8 is fixed to the side of the front-stage compressor disk 2 near the rear-stage compressor disk 3. The central axis of the guide ring 8 is collinear with the central axis of the front-stage compressor disk 2. The outer circumferential surface of the guide ring 8 is arc-shaped, and the arc-shaped surface is concave towards the central axis near the guide ring 8. Multiple guide vanes 81 are fixed to the outer circumferential surface of the guide ring 8, and the multiple guide vanes 81 are equally spaced along the circumference of the guide ring 8.

[0049] When the cold air passes between multiple guide plates 61 and multiple guide plates 62, the cold air is blown onto the outer circumferential surface of the guide ring 8, and then the cold air is divided by multiple guide vanes 81, so that the cold air is evenly distributed circumferentially when passing through the axial air intake section 1.

[0050] The implementation principle of a gas turbine compressor vortex reducer mechanism in this application embodiment is as follows: when cooling gas is introduced into the disk cavity 42 through the drum hole 41, the cold gas in the disk cavity 42 passes through the guide hole 51, and is guided and diverted by multiple guide plates 61 and multiple guide plates 62, so that the cold gas enters the axial air intake section 1, reducing the formation of free vortex structure in the disk wall, thereby reducing the large pressure loss of the cold gas.

[0051] By starting two motors 78, motors 78 drive gear 2 79 to rotate, gear 2 79 drives gear ring 73 to rotate, gear ring 73 simultaneously drives multiple gears 1 72 to rotate, gear 1 72 drives sleeve 71 to rotate, causing sleeve 71 to drive push rod 65 to move along the axial direction of sleeve 71, causing push rod 65 to drive sliding rod 66 to move along the axial direction of sleeve 71, thereby causing multiple guide plates 1 61 and multiple guide plates 2 62 to swing and adjust the tilt angle of multiple guide plates 1 61 and multiple guide plates 2 62, thereby adjusting the flow rate of cold air through axial air intake section 1.

[0052] The above are all preferred embodiments of this application and are not intended to limit the scope of protection of this application. Therefore, all equivalent changes made in accordance with the structure, shape and principle of this application should be covered within the scope of protection of this application.

Claims

1. A vortex suppressor mechanism for a gas turbine compressor, characterized in that: The system includes an axial air intake section (1), a drum (4), and a front-stage compressor disk (2) and a rear-stage compressor disk (3) arranged opposite to each other. The two ends of the drum (4) are respectively fixed to the inner sides of the front-stage compressor disk (2) and the rear-stage compressor disk (3). The front-stage compressor disk (2), the rear-stage compressor disk (3), and the drum (4) form a disk cavity (42). The disk cavity (42) is connected to the axial air intake section (1). The outer circumferential surface of the drum (4) has multiple drum holes (41), each of which is connected to the disk cavity (42). An annular partition (5) is provided between the front stage compressor disk (2) and the rear stage compressor disk (3). The outer circumferential surface of the annular partition (5) is provided with a plurality of drainage holes (51) corresponding to a plurality of drum holes (41). A plurality of sets of flow guiding components (6) are provided between the front stage compressor disk (2) and the rear stage compressor disk (3). The plurality of sets of flow guiding components (6) are located between the central axis of the annular partition (5) and the front stage compressor disk (2). The plurality of sets of flow guiding components (6) correspond to a plurality of drainage holes (51). Each set of flow guiding components (6) includes a plurality of flow guiding plates (61). The compressor disk consists of a front-stage compressor disk (2) and multiple guide vanes (62). Multiple guide vanes (61) and multiple guide vanes (62) are arranged opposite to each other. The ends of the guide vanes (61) and the guide vanes (62) that are far apart from each other are respectively located on the sides of the front-stage compressor disk (2) opposite to the rear-stage compressor disk (3). The guide vanes (61) and the guide vanes (62) are respectively inclined towards the central axis of the front-stage compressor disk (2). The sides of the multiple guide vanes (61) that are far away from the central axis of the front-stage compressor disk (2) and the sides of the multiple guide vanes (62) that are far away from the central axis of the front-stage compressor disk (2) are respectively inclined towards the central axis of the front-stage compressor disk (2). The sides of the central axis of the compressor disk (2) are all provided with moving grooves (63), and push rods (65) are inserted in the multiple moving grooves (63). The opposite inner sides of the moving grooves (63) are all provided with sliding grooves (64). A sliding rod (66) is slidably arranged in the sliding groove (64). One end of the sliding rod (66) near the central axis of the push rod (65) is fixedly connected to the outer peripheral surface of the push rod (65). The sides of the front stage compressor disk (2) and the rear stage compressor disk (3) are respectively provided with driving mechanisms for driving the push rod (65) to move radially along the front stage compressor disk (2).

2. The gas turbine compressor vortex reducer mechanism according to claim 1, characterized in that: The plurality of guide vanes one (61) and the plurality of guide vanes two (62) are all arc-shaped plates. The arc surfaces of the plurality of guide vanes one (61) and the plurality of guide vanes two (62) protrude toward the direction away from the central axis of the front stage compressor disk (2). The area of ​​the plurality of guide vanes one (61) and the area of ​​the plurality of guide vanes two (62) gradually decrease toward the direction closer to the central axis of the front stage compressor disk (2).

3. The gas turbine compressor vortex reducer mechanism according to claim 1, characterized in that: The drive mechanism includes multiple connecting plates (7) fixed on the opposite sides of the front stage compressor disk (2) and the rear stage compressor disk (3). The multiple connecting plates (7) correspond to multiple push rods (65). A sleeve (71) is provided on the connecting plate (7). The sleeve (71) is rotatably connected to the connecting plate (7). The push rod (65) passes through the sleeve (71) and is threadedly connected to the sleeve (71). The opposite sides of the front stage compressor disk (2) and the rear stage compressor disk (3) are respectively provided with drive components for driving the multiple sleeves (71) to rotate.

4. The gas turbine compressor vortex suppressor mechanism according to claim 3, characterized in that: The drive assembly includes a gear 1 (72) disposed on the outer peripheral surface of the sleeve (71). The front stage compressor disk (2) and the rear stage compressor disk (3) are rotatably provided with gear rings (73). The gear rings (73) mesh with multiple gear 1s (72). The front stage compressor disk (2) and the rear stage compressor disk (3) are respectively provided with drive members for driving the two gear rings (73) to rotate.

5. The gas turbine compressor vortex reducer mechanism according to claim 4, characterized in that: The drive unit includes two motors (78), which are respectively fixed on the sides of the front stage compressor disk (2) and the rear stage compressor disk (3). A gear (79) is provided on the output shaft of the motor (78), and the gear (79) meshes with the gear ring (73).

6. The gas turbine compressor vortex reducer mechanism according to claim 4, characterized in that: The front stage compressor disk (2) and the rear stage compressor disk (3) are provided with annular grooves (74) on their opposite sides. The two gear rings (73) are fixed with rings (75) on their opposite sides. The rings (75) are inserted into the annular grooves (74). The inner circumferential surface of the annular grooves (74) is provided with annular grooves (76). The outer circumferential surface of the rings (75) is fixed with rings (77). The rings (77) are disposed in the annular grooves (76).

7. The gas turbine compressor vortex reducer mechanism according to claim 1, characterized in that: A guide ring (8) is fixed on the side of the front stage compressor disk (2) near the rear stage compressor disk (3). The central axis of the guide ring (8) is collinear with the central axis of the front stage compressor disk (2), and the outer circumferential surface of the guide ring (8) is an arc-shaped surface.

8. The gas turbine compressor vortex reducer mechanism according to claim 7, characterized in that: Multiple guide vanes (81) are fixed on the outer circumferential surface of the guide ring (8), and the multiple guide vanes (81) are arranged at equal intervals along the circumference of the guide ring (8).

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