Patent Application
20180100399
The invention relates to a method for profiling a turbine rotor blade for an axial flow machine.
The trend in the design of blades for an axial flow machine is toward increasing the aspect ratio of the blades and making the blades thinner. The blades designed in such a way tend to flutter during the operation of the axial flow machine. The fluttering is a self-induced vibration at the natural frequency of the blade. This vibration may be a longitudinal vibration of the blade with a vibration node at the root of the blade. Energy is thereby transferred from the fluid flowing in the axial flow machine to the blade. With repeated load changes of the axial flow machine, the fluttering may lead to material fatigue of the blade (high cycle fatigue). The material fatigue may lead to the formation of a crack and necessitate a cost-intensive replacement of the blade.
Fluttering is conventionally prevented by reducing the load acting on the blade. This however disadvantageously leads to a reduction in the efficiency of the axial flow machine. Furthermore, damping elements are conventionally provided, such as for example a shroud, which damps the fluttering of the blades. This however is a structurally complex solution. It would therefore be desirable to design the blade in such a way that it does not tend to flutter during the operation of the axial flow machine.
The object of the invention is to provide a method for profiling a blade for an axial flow machine in which the blade tends less to flutter.
The method according to the invention for profiling a turbine rotor blade for an axial flow machine has the steps of: providing a geometrical model of a blade profile, which has a mean camber line of a profile section of the turbine rotor blade; determining boundary conditions for a flow flowing around the turbine rotor blade; changing the mean camber line in such a way that the flow that is established by the boundary conditions produces the maximum of the difference of the isentropic Mach number between the pressure side and the suction side of the turbine rotor blade in a blade portion that extends from the blade trailing edge in the direction of the blade leading edge and the length of which is 65% of the length S of the blade chord. The mean camber line is the line of the profile section defined by points at the same distance from the pressure side as from the suction side. The blade chord denotes the path in the profile section from the blade leading edge to the blade trailing edge. Calculations have shown that, if the maximum of the difference of the isentropic Mach number is arranged in the blade portion according to the invention, the unstable pressure distribution changes in such a way that to the greatest extent local damping regions and local exciting regions compensate for one another. As a result, the blades designed in such a way tend much less to flutter than conventionally designed blades. The low tendency to flutter allows the blades to be subjected to greater loading than the conventionally designed blades. Moreover, there is advantageously no need for additional damping elements, such as for example a shroud, to be provided.
The mean camber line is formed by a first fourth-degree polynomial, which describes the mean camber line from the blade leading edge to an extreme point, and a second fourth-degree polynomial, which describes the mean camber line from the extreme point to the blade trailing edge, the extreme point being the point of the mean camber line that is at the maximum distance from the blade chord. The distance denotes the length of a path extending at right angles from the blade chord to the mean camber line. It is advantageous that the first polynomial is formed by using a leading-edge mean camber-line angle, which is the angle between the leading-edge tangent of the mean camber line and the blade chord, the length xS1 from the blade leading edge to the point of the blade chord that is at the maximum distance from the mean camber line, and the length S1, which is the distance from the extreme point to the blade chord, the second polynomial being formed by using a trailing-edge mean camber-line angle, which is the angle between the trailing-edge tangent of the mean camber line and the blade chord, the length S-xS1 from the blade trailing edge to the point of the blade chord that is at the maximum distance from the mean camber line, and the length S2, which is the distance from the mean camber line to the point of the blade chord that is at the distance xS1+0.5*(S-xS1) from the blade trailing edge, where S is the length of the blade chord. If a slope of zero is assumed for the extreme point, the first polynomial and the second polynomial are sufficiently determined by these parameters.
It is advantageous that the mean camber line is changed in such a way that S1 is from 10.3% to 11.3% of the length S, xS1 is from 35.1% to 38.4% of the length S of the blade chord, S2 is from 64.8% to 67.9% of the length S1, the trailing-edge mean camber-line angle is from 15.192° to 19.020° and the leading-edge mean camber-line angle is from 37.663° to 39.256°. It is advantageously ensured by these parameters that the blade has only a low tendency to flutter. The mean camber line is advantageously changed in such a way that S1 is 10.8% of the length S, xS1 is 36.8% of the length S, S2 is 66.3% of the length S1, the leading-edge mean camber-line angle is 17.106° and the trailing-edge mean camber-line angle is 38.460°. It is advantageously achieved by these parameters that the blade has a particularly low tendency to flutter.
It is alternatively advantageous that the turbine rotor blade has a transonic portion and the mean camber line in the transonic portion is changed in such a way that S1 is from 7.6874% to 7.9% of the length S, xS1 is from 35.4311% to 36.2% of the length S, S2 is from 63% to 65% of the length S1, the trailing-edge mean camber-line angle is from 11.0° to 12.3° and the leading-edge mean camber-line angle is from 29.0° to 31.0°. These parameters have the effect that a compression shock occurring during the operation of the axial flow machine under the boundary conditions occurs a long way downstream and with a low Mach number gradient. A fluttering turbine rotor blade causes disturbances in the flow. These disturbances may change the position of the compression shock that occurs at an adjacent turbine rotor blade. However, because the compression shock is arranged a long way downstream, the disturbances can only change the position of the compression shock to a small extent. As a result, a fluttering turbine rotor blade can only induce the fluttering of an adjacent turbine rotor blade to a small degree, as a result of which the overall fluttering tendency is low. In addition, the low Mach number gradient for the compression shock means that fluttering induced by the compression shock is advantageously reduced.
It is advantageous that the turbine rotor blade is free-standing. This means that no damping elements, such as for example a shroud, are provided.
It is advantageous that the geometrical model has a thickness that varies along the mean camber line, which is left the same during the changing of the mean camber line. Advantageously, here only the mean camber line is changed to reduce the tendency of the blade to flutter, which is advantageously a simple method with only few parameters to be changed.
It is advantageous that the boundary conditions of the flow are obtained from the nominal operating condition of the axial flow machine. It is also advantageous that it is a steady-state flow. The isentropic Mach numbers are advantageously determined experimentally and/or determined computationally. It is advantageous that the method is repeated for different profile sections of the turbine rotor blade. As a result, a design of the turbine rotor blade along its height takes place. The profile section advantageously lies on a cylinder surface or a cone surface of which the axes coincide with the axis of the axial flow machine, on an S1 flow surface or in a tangential plane of the axial flow machine.
The axial flow machine is advantageously a gas turbine or a steam turbine. The method is advantageously carried out for profile sections that lie in the radially outer half of the turbine rotor blade; in particular, the method is only carried out for the profile sections that lie in the radially outer half of the turbine rotor blade.
The turbine rotor blade according to the invention for an axial flow machine has a blade profile that has a mean camber line of a profile section of the turbine rotor blade, the mean camber line being formed in such a way that, on the basis of boundary conditions for a flow flowing around the turbine rotor blade, the flow that is established produces the maximum of the difference of the isentropic Mach number between the pressure side and the suction side of the turbine rotor blade in a blade portion that extends from the blade trailing edge in the direction of the blade leading edge and the length of which is 65% of the length S of the blade chord.
It is advantageous that the mean camber line is formed by a first fourth-degree polynomial, which describes the mean camber line from the blade leading edge to an extreme point, and a second fourth-degree polynomial, which describes the mean camber line from the extreme point to the blade trailing edge, the extreme point being the point of the mean camber line that is at the maximum distance from the blade chord, the first polynomial being formed by using a leading-edge mean camber-line angle, which is the angle between the leading-edge tangent of the mean camber line and the blade chord, the length xS1 from the blade leading edge to the point of the blade chord that is at the maximum distance from the mean camber line, and the length S1, which is the distance from the extreme point to the blade chord, the second polynomial being formed by using a trailing-edge mean camber-line angle, which is the angle between the trailing-edge tangent of the mean camber line and the blade chord, the length S-xS1 from the blade trailing edge to the point of the blade chord that is at the maximum distance from the mean camber line, and the length S2, which is the distance from the mean camber line to the point of the blade chord that is at the distance xS1+0.5*(S-xS1) from the blade trailing edge, where S is the length of the blade chord.
It is advantageous that the mean camber line is made such that S1 is from 10.3% to 11.3% of the length S, xS1 is from 35.1% to 38.4% of the length S, S2 is from 64.8% to 67.9% of the length S1, the trailing-edge mean camber-line angle is from 15.192° to 19.020° and the leading-edge mean camber-line angle is from 37.663° to 39.256°. Alternatively, it is advantageous that the turbine rotor blade has a transonic portion and the mean camber line in the transonic portion is made such that S1 is from 7.6874% to 7.9% of the length S, xS1 is from 35.4311% to 36.2% of the length S, S2 is from 63% to 65% of the length S1, the trailing-edge mean camber-line angle is from 11.0° to 12.3° and the leading-edge mean camber-line angle is from 29.0° to 31.0°.
The axial flow machine according to the invention has a turbine rotor blade according to the invention, the turbine rotor blade being free-standing and the axial flow machine being in particular a gas turbine or a steam turbine.
The invention is explained in more detail below on the basis of the accompanying schematic drawings, in which:
As can be seen from
The mean camber line 3 is formed by a first fourth-degree polynomial 11 and a second fourth-degree polynomial 12. The first polynomial 11 describes the mean camber line 3 from the blade leading edge 4 to an extreme point 30. The extreme point 30 is the point of the mean camber line 3 that is at the maximum distance from the blade chord 13. The second polynomial 12 describes the mean camber line 3 from the extreme point 30 to the blade trailing edge 5. Likewise depicted in
The first polynomial 11 is formed by choosing the leading-edge mean camber-line angle LESA, the length xS1 from the blade leading edge 4 to the point (xS1,0) on the blade chord 13 that is at the maximum distance from the mean camber line 13, and the length S1, which is the distance from the point (xS1,0) to the extreme point 30. The fact that the slope of the extreme point 30 is zero and the blade leading edge 4 lies at the origin of the system of coordinates means that the first polynomial 11 is sufficiently determined. The second polynomial 12 is formed by choosing the trailing-edge mean camber-line angle TESA, the length S-xS1 from the blade trailing edge 5 to the point (xS1,0) on the blade chord 13, and the length S2, which is the distance from the point (xS1+0.5*(S-xS1),0) to the mean camber line 3. The fact that the slope of the extreme point 30 is zero and the blade trailing edge 5 lies at the point (S,0) means that the second polynomial 12 is sufficiently determined.
In the method for profiling the blade, the geometrical model of the blade profile is provided in the way described for
In order to achieve the effect that the maximum of the difference of the isentropic Mach number is in the blade portion according to the invention, the parameters describing the first polynomial 11 and the second polynomial 12 may assume for example the following values:
The Mach number variations 22 to 25 show that, for the conventionally designed turbine rotor blade, the difference of the Mach number variations 25 and 23 is greater in the front region of the blade 14 than in the rear region of the turbine rotor blade 14. By contrast, the difference of the Mach number variations 24 and 22 for the blade 15 profiled according to the invention is greater in the rear region of the turbine rotor blade 15 than in the front region of the turbine rotor blade 15. The maximum of the difference of the turbine rotor blade 15 designed according to the invention is located substantially at a length of the blade chord 13 of 0.5*S.
In order to achieve the effect that the maximum of the difference of the isentropic Mach number is in the blade portion according to the invention, in the case of an alternative turbine rotor blade the parameters describing the first polynomial 11 and the second polynomial 12 may alternatively assume for example the following values in a transonic portion of a turbine rotor blade:
d(t)=a0·tFSE+a1·t+a2·t2+a3·t3,
where t goes from 0 to 1, the blade leading edge 4 lying at 0 and the blade trailing edge lying at 1. The polynomial is formed by choosing the leading-edge radius of curvature RLE, the length xD1 from the blade leading edge 4 to the point (xD1,0) on the blade chord 13, at which there is the maximum thickness D1 of the alternative turbine rotor blade, the thickness d2, which is the thickness of the alternative turbine rotor blade at the point (xD1+0.5*(S-xD1),0), and the trailing-edge wedge angle TEWA. The blade also has at the blade trailing edge 5 a portion tapering to a point toward the blade trailing edge 5, which starts from a thickness d3 and falls to zero. The thickness d3 may be in a range from 96% to 99.9% of S.
The aforementioned variables may assume the following values:
Although the invention has been more specifically illustrated and described in detail by the preferred exemplary embodiment, the invention is not restricted by the disclosed examples and other variations can be derived herefrom by a person skilled in the art without departing from the scope of protection of the invention.
| Number | Date | Country | Kind |
|---|---|---|---|
| 15165330.0 | Apr 2015 | EP | regional |
This application is the US National Stage of International Application No. PCT/EP2016/058559 filed Apr. 18, 2016, and claims the benefit thereof. The International Application claims the benefit of European Application No. EP15165330 filed Apr. 28, 2015. All of the applications are incorporated by reference herein in their entirety.
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/EP2016/058559 | 4/18/2016 | WO | 00 |
