A rotor, for a compressor of a gas turbine, comprising a rotatable support for rotation about an axis of rotation and a plurality of blades. Each blade comprising a hub, a leading edge and a trailing edge and a chord is defined between the leading edge and the trailing edge. Each of the blades exten
A rotor, for a compressor of a gas turbine, comprising a rotatable support for rotation about an axis of rotation and a plurality of blades. Each blade comprising a hub, a leading edge and a trailing edge and a chord is defined between the leading edge and the trailing edge. Each of the blades extends from its hub away from the rotatable support and at least one of the blades has a hub-thickness to chord ratio greater than 10 percent. The leading edge of the at least one of the blades at its hub is positioned at a leading-edge-hub-radius from a position of the axis of rotation and the trailing edge of the at least one of the blades is positioned at its hub at a trailing-edge-hub-radius from the position of the axis of rotation. The trailing-edge-hub-radius is greater than the leading-edge-hub-radius.
대표청구항▼
1. A rotor, for a compressor of a gas turbine, comprising: a rotatable support for rotation about an axis of rotation; anda plurality of blades, each blade comprising a hub, a leading edge and a trailing edge wherein a chord is defined between the leading edge and the trailing edge;wherein: each of
1. A rotor, for a compressor of a gas turbine, comprising: a rotatable support for rotation about an axis of rotation; anda plurality of blades, each blade comprising a hub, a leading edge and a trailing edge wherein a chord is defined between the leading edge and the trailing edge;wherein: each of the blades extends from its hub away from the rotatable support;at least one of the blades has a hub-thickness to chord ratio greater than 10 percent;the leading edge of the at least one of the blades at its hub is positioned at a leading-edge-hub-radius from a position of the axis of rotation;the trailing edge of the at least one of the blades is positioned at its hub at a trailing-edge-hub-radius from the position of the axis of rotation;the trailing-edge-hub-radius is greater than the leading-edge-hub-radius;a distance A1 is defined as the distance between lines projected from the leading edges of two adjacent blades and a distance A2 is defined as the distance between lines projected from the trailing edges of the two adjacent blades, the ratio A2/A1 being greater than 2.2;an axial chord of the at least one of the blades is defined by a projection of the chord onto the axis of rotation and has an axial chord length; andfor the at least one of the blades the difference between the trailing-edge-hub-radius and the leading-edge-hub-radius is greater than 25 percent of the axial chord length. 2. A rotor as claimed in claim 1, wherein the rotatable support is a disk and the plurality of blades and the disk are one integral part. 3. A rotor as claimed in claim 1, wherein the rotor is a gas turbine front-stage compressor rotor. 4. A rotor as claimed in claim 1, wherein the rotor is a gas turbine high pressure core-compressor rotor. 5. A rotor as claimed in claim 1, wherein the at least one of the blades has a hub-thickness to chord ratio greater than 12 percent. 6. A rotor as claimed in claim 5, wherein: an angle X1 is defined between a line projected out from the leading edge of one of the blades and a projection of the direction of the axis of rotation;an angle X2 is defined between a line projected out from the trailing edge of the said one of the blades and a projection of the direction of the axis of rotation; anda camber of the blade given by X1-X2 is greater than 38 degrees. 7. A rotor as claimed claim 1, wherein the at least one of the blades has a hub-thickness to chord ratio greater than 13 percent. 8. A rotor as claimed in claim 1, wherein the position of the hub of the at least one of the blades from the position of the axis of rotation is defined by a hub radius of the at least one of the blades and the hub radius has a profile between the leading edge of the at least one of the blades and the trailing edge of the at least one of the blades and wherein there is a single point of inflection in the profile of the hub radius between the leading edge of the at least one of the blades and the trailing edge of the at least one of the blades. 9. A rotor as claimed in claim 8, wherein the profile of the hub radius between the leading edge of the at least one of the blades and the trailing edge of the at least one of the blades is axi-symmetric with regard to the axis of rotation. 10. A gas turbine comprising a rotor as claimed in claim 1. 11. A rotor as claimed in claim 1, wherein: an angle X1 is defined between a line projected out from the leading edge of one of the blades and a projection of the direction of the axis of rotation;an angle X2 is defined between a line projected out from the trailing edge of the said one of the blades and a projection of the direction of the axis of rotation; anda camber of the blade given by X1-X2 is greater than 38 degrees. 12. A rotor as claimed in claim 1, wherein the rotor has a de Haller number of approximately 0.7 along a full extent of the blade. 13. A rotor, for a compressor of a gas turbine, comprising: a rotatable support for rotation about an axis of rotation; anda plurality of blades, each blade comprising a hub, a leading edge and a trailing edge wherein a chord is defined between the leading edge and the trailing edge and wherein:each of the blades extends from its hub away from the rotatable support;at least one of the blades has a hub-thickness to chord ratio greater than 10 percent;the leading edge of the at least one of the blades at its hub is positioned at a leading-edge-hub-radius from a position of the axis of rotation; andthe trailing edge of the at least one of the blades is positioned at its hub at a trailing-edge-hub-radius from the position of the axis of rotation wherein the trailing-edge-hub-radius is greater than the leading-edge-hub-radius;wherein an axial chord of the at least one of the blades is defined by a projection of the chord onto the axis of rotation and has an axial chord length, and wherein for the at least one of the blades the difference between the trailing-edge-hub-radius and the leading-edge-hub-radius is greater than 20 percent of the axial chord length. 14. A rotor as claimed in claim 13, wherein the rotor has a de Haller number of approximately 0.7 along a full extent of the blade. 15. A rotor, for a compressor of a gas turbine, comprising: a rotatable support for rotation about an axis of rotation; anda plurality of blades, each blade comprising a hub, a leading edge and a trailing edge wherein a chord is defined between the leading edge and the trailing edge;wherein: each of the blades extends from its hub away from the rotatable support;at least one of the blades has a hub-thickness to chord ratio greater than 10 percent;the leading edge of the at least one of the blades at its hub is positioned at a leading-edge-hub-radius from a position of the axis of rotation;the trailing edge of the at least one of the blades is positioned at its hub at a trailing-edge-hub-radius from the position of the axis of rotation;the trailing-edge-hub-radius is greater than the leading-edge-hub-radius;an angle X1 is defined between a line projected out from the leading edge of one of the blades and a projection of the direction of the axis of rotation;an angle X2 is defined between a line projected out from the trailing edge of the said one of the blades and a projection of the direction of the axis of rotation;a camber of the blade given by X1-X2 is greater than 38 degrees;an axial chord of the at least one of the blades is defined by a projection of the chord onto the axis of rotation and has an axial chord length; andfor the at least one of the blades the difference between the trailing-edge-hub-radius and the leading-edge-hub-radius is greater than 25 percent of the axial chord length. 16. A rotor according to claim 15, wherein the hub-thickness to chord ratio is greater than 12 percent. 17. A rotor as claimed in claim 15, wherein the rotatable support is a disk and the plurality of blades and the disk are one integral part. 18. A rotor as claimed claim 15, wherein the at least one of the blades has a hub-thickness to chord ratio greater than 13 percent. 19. A rotor as claimed in claim 15, wherein: the position of the hub of the at least one of the blades from the position of the axis of rotation is defined by a hub radius of the at least one of the blades;the hub radius has a profile between the leading edge of the at least one of the blades and the trailing edge of the at least one of the blades; andthere is a single point of inflection in the profile of the hub radius between the leading edge of the at least one of the blades and the trailing edge of the at least one of the blades. 20. A rotor as claimed in claim 15, wherein the profile of the hub radius between the leading edge of the at least one of the blades and the trailing edge of the at least one of the blades is axi-symmetric with regard to the axis of rotation. 21. A rotor as claimed in claim 15, wherein the rotor has a de Haller number of approximately 0.7 along a full extent of the blade.
연구과제 타임라인
LOADING...
LOADING...
LOADING...
LOADING...
LOADING...
이 특허에 인용된 특허 (1)
Matheny, Alfred P.; Kite, Edwin L., Replaceable leading edge insert for an IBR.
※ AI-Helper는 부적절한 답변을 할 수 있습니다.