CN214366235U - Groove structure for reducing secondary flow loss of gas turbine end wall - Google Patents
Groove structure for reducing secondary flow loss of gas turbine end wall Download PDFInfo
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- CN214366235U CN214366235U CN202022705750.2U CN202022705750U CN214366235U CN 214366235 U CN214366235 U CN 214366235U CN 202022705750 U CN202022705750 U CN 202022705750U CN 214366235 U CN214366235 U CN 214366235U
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Abstract
The utility model discloses a reduce groove structure of gas turbine end wall secondary flow loss, this groove structure set up at the turbine channel end wall, and groove structure's cross section is symmetrical pattern, and its symmetry axis is unanimous with the streamline direction, and groove structure comprises a plurality of geometric equivalence scale recesses, and the axial position and the adjacent interval of each scale recess are the same. The utility model discloses based on groove structure reduces the flow resistance, adsorbs the mechanism that flows near the wall, through reducing turbine channel end wall transverse pressure gradient, reduce the interior horse shoe vortex of passageway and passageway vortex intensity, reduce the secondary flow loss, improve turbine wholeness ability.
Description
Technical Field
The utility model belongs to gas turbine axial compressor turbine field, concretely relates to reduce groove structure of gas turbine end wall secondary flow loss.
Background
The gas turbine is widely applied to the aspects of aviation, ship power, power generation and the like as a power source. The gas turbine working under severe conditions of high pressure, high temperature, high speed and the like is one of key parts of the gas turbine, and along with the continuous improvement of the performance of the modern gas turbine, higher requirements are put forward on the design of the gas turbine. The flow field inside the turbine channel has a strong three-dimensional shape and unsteadiness, and the complicated flow causes various losses. Wherein, the secondary flow loss can reach 30-50% of the total flow loss, and is an important component of the flow loss. Therefore, it is important to develop new measures for reducing the secondary flow loss and to provide a method for reducing the secondary flow loss in engineering practice.
In a turbine channel, the secondary flow phenomenon is generally represented by various vortex system structures, such as horseshoe vortex, channel vortex and the like, and is mainly concentrated in an end region, and the generation and development of the vortex systems are not only related to the geometric parameters of a turbine blade cascade, but also influenced by the pneumatic parameters of incoming flow conditions, flow field quality and the like. According to the existing research, the methods for reducing the secondary flow include changing the shape of a leading edge, an end wall vane, an asymmetric end wall and the like, and the methods are all used for reducing the loss by changing the vortex structure of the end region.
SUMMERY OF THE UTILITY MODEL
An object of the utility model is to provide a reduce groove structure of gas turbine end wall secondary flow loss, reduce the flow resistance based on groove structure, adsorb the mechanism that flows near the wall, through reducing turbine channel end wall transverse pressure gradient, reduce interior water chestnut whirlpool of passageway and passageway vortex intensity, reduce the secondary flow loss, improve turbine wholeness ability.
The utility model discloses a following technical scheme realizes:
a groove structure for reducing the secondary flow loss of the end wall of a turbine of a gas turbine is arranged on the end wall of a turbine channel, the cross section of the groove structure is a symmetrical graph, the symmetrical axis of the groove structure is consistent with the direction of a streamline, the groove structure is composed of a plurality of microscale grooves with the same geometry, and the axial position and the adjacent distance of each microscale groove are the same.
The utility model discloses a further improvement lies in, groove structure's cross section is rectangle, triangle-shaped or semi-circular.
The utility model discloses further improvement lies in, groove structure is apart from turbine blade leading edge distance for S at the starting point on the turbine passageway end wall, and the termination point is apart from turbine blade trailing edge distance for T, and wherein, the value scope of S is 10% C ~ 50% C, and the value scope of T is 10% C ~ 40% C, and C shows blade axial chord length.
The utility model discloses further improvement lies in, groove structure leftmost distance with the blade suction surface is L, and L's value range is 2% D ~ 5% D, and D represents the cascade bars apart from.
The utility model discloses further improvement lies in, and the width of each microscale recess of groove structure is W, and W's scope is 0.2% D ~ 2% D, and D represents the cascade grid distance.
The utility model discloses further improvement lies in, and the interval between each microscale recess of groove structure is G, and G's scope is 0.1% D ~ 1% D.
The utility model discloses further improvement lies in, the degree of depth of each microscale recess of groove structure is H, and H's scope is 0.2% D ~ 2% D.
The utility model discloses further improvement lies in, groove structure's microscale recess's quantity is N, and N's scope is 10 ~ 30.
The utility model discloses at least, following profitable technological effect has:
the utility model provides a reduce groove structure of gas turbine end wall secondary flow loss takes as above at the turbine passageway end wall the groove treatment after, the turbine passageway end wall transverse pressure gradient reduces to some extent, has weakened the development of horse hoof whirlpool and passageway whirlpool to secondary flow loss has been reduced. Therefore, the utility model discloses be different from other methods in the past, the structural style that this method adopted is simple, the effect is obvious, implement the convenience, simultaneously because the yardstick is less, and the pneumatic loss that additionally brings is less.
Drawings
FIG. 1 is a schematic illustration of near wall vortices in a turbine passageway.
Fig. 2 is a three-dimensional schematic diagram of the groove structure of the turbine channel end wall of the present invention.
Fig. 3 is a top view of the present invention.
Fig. 4 is a cross-sectional view of the present invention.
Fig. 5 is a schematic cross-sectional view of the rectangular micro-scale groove of the present invention.
Fig. 6 is a schematic cross-sectional view of the triangular micro-scale groove of the present invention.
Fig. 7 is a schematic cross-sectional view of the semicircular micro-scale groove of the present invention.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more clearly understood, the present invention will be further described in detail with reference to the accompanying drawings and examples.
As shown in fig. 1 to 7, the utility model provides a pair of reduce groove structure of gas turbine end wall secondary flow loss through adding groove structure at turbine cascade passageway end wall, reduces turbine passageway end wall transverse pressure gradient, reduces the interior horseshoe vortex of passageway and passageway vortex intensity, reduces the secondary flow loss. The cross section of the groove structure is a symmetrical graph, the symmetry axis of the groove structure is consistent with the direction of the streamline, the groove structure is composed of a plurality of microscale grooves with the same geometry, and the axial position and the adjacent distance of each microscale groove are the same.
The axial chord length of the turbine blade is C, and the grid pitch of the blade grid is D. In the present embodiment, the axial chord length C of the blade is taken as 100mm, and the cascade pitch D is taken as 80 mm.
In the present embodiment, the cross section of the added groove structure is rectangular.
The distance S from the starting point of the added groove structure to the leading edge of the turbine blade is in the range of 5% C to 30% C, and in the present embodiment, S is preferably 20% C, i.e., 20 mm.
The distance T of the added groove structure termination point from the turbine blade trailing edge ranges from 10% C to 40% C, in this embodiment T is preferably 15% C, i.e. 15 mm.
The distance L between the leftmost side of the added groove structure and the suction surface of the turbine blade ranges from 2% D to 5% D, and in the embodiment, L is preferably 2% D, namely 1.6 mm.
The width W of each micro-scale groove of the added groove structure ranges from 0.2% D to 2% D, and in the present embodiment, W is preferably 1% D, i.e., 0.8 mm.
The pitch G between the individual micro-scale grooves of the added groove structure ranges from 0.1% D to 1% D, in this embodiment H is preferably 0.8% D, i.e. 0.64 mm.
The depth H of each micro-scale groove of the added groove structure ranges from 0.2% D to 2% D, in this embodiment, H is preferably 1% D, i.e. 0.8 mm.
The number N of the added micro-scale grooves of the groove structure ranges from 10 to 30, and in the embodiment, N is preferably 10.
In the present embodiment, after the end wall of the turbine passage has the groove structure as described above, the transverse pressure gradient of the end wall of the turbine passage is reduced, the horseshoe vortex and the strength of the channel vortex in the passage are suppressed, the loss of the secondary flow is reduced, and the overall performance of the turbine is improved.
The above description is only for the preferred embodiment of the present invention, and is not intended to limit the present invention, and any modifications, equivalent replacements, improvements, etc. made within the spirit and principle of the present invention should be included within the scope of the present invention.
Claims (8)
1. A groove structure for reducing the loss of secondary flow of a turbine end wall of a gas turbine is characterized in that the groove structure is arranged on the end wall of a turbine channel, the cross section of the groove structure is a symmetrical graph, the symmetrical axis of the groove structure is consistent with the streamline direction, the groove structure is composed of a plurality of micro-scale grooves with the same geometry, and the axial position and the adjacent distance of each micro-scale groove are the same.
2. A groove structure for reducing the loss of secondary flow at the end wall of a gas turbine according to claim 1, wherein the cross section of the groove structure is rectangular, triangular or semicircular.
3. The groove structure for reducing the secondary flow loss of the end wall of the gas turbine as claimed in claim 1, wherein the starting point of the groove structure on the end wall of the turbine passage is located at a distance S from the leading edge of the turbine blade, and the ending point of the groove structure is located at a distance T from the trailing edge of the turbine blade, wherein S is 10% C to 50% C, T is 10% C to 40% C, and C represents the axial chord length of the blade.
4. The groove structure for reducing the loss of the secondary flow of the end wall of the gas turbine according to claim 1, wherein the distance between the leftmost side of the groove structure and the suction surface of the blade is L, the value of L ranges from 2% D to 5% D, and D represents the grid pitch of the blade.
5. The groove structure for reducing the secondary flow loss of the end wall of the gas turbine according to claim 1, wherein each micro-scale groove of the groove structure has a width W, W is in a range of 0.2% D to 2% D, and D represents a blade row pitch.
6. A groove structure for reducing the loss of secondary flow at the end wall of a gas turbine according to claim 1, wherein the distance between the micro-scale grooves of the groove structure is G, and G is in the range of 0.1% D to 1% D.
7. A groove structure for reducing the loss of secondary flow at the end wall of a gas turbine according to claim 1, wherein each micro-scale groove of the groove structure has a depth H, and H is in the range of 0.2% D to 2% D.
8. The groove structure for reducing the secondary flow loss of the end wall of the gas turbine according to claim 1, wherein the number of the micro-scale grooves of the groove structure is N, and N is in the range of 10-30.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202022705750.2U CN214366235U (en) | 2020-11-20 | 2020-11-20 | Groove structure for reducing secondary flow loss of gas turbine end wall |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202022705750.2U CN214366235U (en) | 2020-11-20 | 2020-11-20 | Groove structure for reducing secondary flow loss of gas turbine end wall |
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| CN214366235U true CN214366235U (en) | 2021-10-08 |
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| CN202022705750.2U Active CN214366235U (en) | 2020-11-20 | 2020-11-20 | Groove structure for reducing secondary flow loss of gas turbine end wall |
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