IPC분류정보
국가/구분 |
United States(US) Patent
등록
|
국제특허분류(IPC7판) |
|
출원번호 |
US-0870437
(2004-06-18)
|
등록번호 |
US-7484937
(2009-02-03)
|
우선권정보 |
EP-04090215(2004-06-02) |
발명자
/ 주소 |
|
출원인 / 주소 |
- Rolls Royce Deutschland Ltd & Co KG
|
대리인 / 주소 |
|
인용정보 |
피인용 횟수 :
7 인용 특허 :
13 |
초록
▼
The compressor blades of an aircraft engine are, in at least one natural-vibration critical area, designed such that at the blade leading edge (6), the leading edge shock wave (14) attaches to the leading edge (6), as a result of which a laminar boundary layer flow (7) on the suction side (13) quick
The compressor blades of an aircraft engine are, in at least one natural-vibration critical area, designed such that at the blade leading edge (6), the leading edge shock wave (14) attaches to the leading edge (6), as a result of which a laminar boundary layer flow (7) on the suction side (13) quickly transitions into a turbulent boundary layer flow (9) which is kept constant and prevented from re-lamination by the further, continuous curvature of the suction side. Therefore, the transition, whose periodic movement is also suppressed, cannot communicate with a suction-side compression shock (10), preventing the compression shock from augmenting the natural vibrations of the blade occurring under certain flight conditions. The blade leading edge can, for example, be designed as an ellipse with a semi-axis ratio equal to or smaller than 1:4.
대표청구항
▼
What is claimed is: 1. A compressor blade for a gas turbine, comprising a suction and a pressure side and a blade leading and trailing edge the blade having a relatively long chord length, wherein the blade leading edge, at a blade tip area, operates at transonic and supersonic velocities, and, in
What is claimed is: 1. A compressor blade for a gas turbine, comprising a suction and a pressure side and a blade leading and trailing edge the blade having a relatively long chord length, wherein the blade leading edge, at a blade tip area, operates at transonic and supersonic velocities, and, in at least one natural-vibration critical area, is structured such that a leading edge shock wave attaches to the blade leading edge, as a result of which an initially laminar boundary layer flow changes, at a transition point a short distance from the blade leading edge into a turbulent boundary layer flow which neither accelerates nor re-laminates within a continuous curvature extending from the blade leading edge on the suction side, thus suppressing a periodic movement of the transition point and preventing the transition point from communicating with a compression shock on the suction side; wherein the blade leading edge is structured as one of an elliptic or parabolic cross-sectional profile having an ellipse ratio (a:b) being equal to or smaller than 1 to 4 to effect attachment of the leading edge shock wave immediately to the blade leading edge. 2. A compressor blade in accordance with claim 1, wherein the blade leading edge includes an area having a serration to effect attachment of the leading edge shock wave. 3. A compressor blade in accordance with claim 1, wherein the blade leading edge includes an area having pocket-style recesses to effect attachment of the leading edge shock wave. 4. A compressor blade in accordance with claim 1, wherein the blade leading edge includes a recessed area acting as a sweep-back to effect attachment of the leading edge shock wave. 5. A compressor blade in accordance with claim 4, wherein the recessed area includes a concave curvature. 6. A compressor blade in accordance with claim 1, wherein the suction side is designed as a reflex which includes a depression in an area of the laminar boundary layer flow on the suction side in order to effect, by pressure increase, a transition into the turbulent boundary layer flow and keep acceleration of the boundary layer flow low and the boundary layer flow itself constant. 7. A compressor blade for a gas turbine compressor, comprising a leading edge, which at a blade tip area, operates at transonic and supersonic velocities, and a suction side which includes a concave depression near the leading edge to cause transition of a boundary layer flow from laminar to turbulent, and keep acceleration of the boundary layer flow low and flow of the turbulent boundary layer constant; the depression being positioned on the suction side between the leading edge and a midpoint of the blade, the depression being shallow with gradually tapering leading and trailing edges and extending across a substantial portion of the distance from the leading edge to the blade midpoint, the depression positioned to begin between the leading edge and a midpoint between the leading edge and the blade midpoint and having a width several times its depth, the blade having a relatively long chord length. 8. A compressor blade for a gas turbine, comprising a leading edge and a suction side, the blade having a relatively long chord length, the blade leading edge, at a blade tip area, operating at transonic and supersonic velocities, and the blade leading edge including an area having pocket-style recesses to attach a leading edge shock wave to the leading edge, and to transition a suction side boundary layer flow from laminar to turbulent at a point a short distance from the leading edge, with the turbulent boundary layer flow neither accelerating nor re-laminating within a continuous curvature extending from the leading edge, thus suppressing a periodic movement of the transition point and preventing the transition point from communicating with a suction side compression shock, the pocket style recesses positioned so that a major portion thereof is positioned on the suction side of the blade adjacent the leading edge. 9. A compressor blade in accordance with claim 8, wherein the suction side is designed as a reflex which includes a depression in an area of the laminar boundary layer flow on the suction side in order to effect, by pressure increase, a transition into the turbulent boundary layer flow and keep acceleration of the boundary layer flow low and the boundary layer flow itself constant. 10. A compressor blade for a gas turbine, comprising a suction and a pressure side and a blade leading and trailing edge, the blade having a relatively long chord length, wherein the blade leading edge, at a blade tip area, operates at transonic and supersonic velocities, and, in at least one natural-vibration critical area, is structured such that a leading edge shock wave attaches to the blade leading edge, as a result of which an initially laminar boundary layer flow changes, at a transition point a short distance from the blade leading edge into a turbulent boundary layer flow which neither accelerates nor re-laminates within a continuous curvature extending from the blade leading edge on the suction side, thus suppressing a periodic movement of the transition point and preventing the transition point from communicating with a compression shock on the suction side; wherein the blade leading edge includes a concavely curved recessed area acting as a sweep-back to effect attachment of the leading edge shock wave, the recessed area starting at a position radially inward from the blade tip and ending before a blade midpoint. 11. A compressor blade in accordance with claim 10, wherein the suction side is designed as a reflex which includes a depression in an area of the laminar boundary layer flow on the suction side in order to effect, by pressure increase, a transition into the turbulent boundary layer flow and keep acceleration of the boundary layer flow low and the boundary layer flow itself constant.
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