Patent Application
20080145228
The present invention relates generally to an array of flow directing elements for a turbomachine and, more particularly, to a rotor blade array configured to interrupt a shock field downstream of rotor blades in the array and reduce shock induced flutter in the rotor blades.
Turbomachinery devices, such as gas turbine engines and steam turbines, operate by exchanging energy with a working fluid using alternating rows of rotating blades and non-rotating vanes. Each blade and vane has an airfoil portion that interacts with the working fluid.
Airfoils have natural vibration modes of increasing frequency and complexity of the mode shape. The simplest and lowest frequency modes are typically referred to as first bending, second bending, and first torsion. First bending is a motion normal to the flat surface of an airfoil in which the entire span of the airfoil moves in the same direction. Second bending is similar to first bending, but with a change in the sense of the motion somewhere along the span of the airfoil, so that the upper and lower portions of the airfoil move in opposite directions. First torsion is a twisting motion around an elastic axis, which is parallel to the span of the airfoil, in which the entire span of the airfoil, on each side of the elastic axis, moves in the same direction.
It is known that turbomachinery blades are subject to destructive vibrations due to unsteady interaction of the blades with the working fluid. One type of vibration is known as flutter, which is an aero-elastic instability resulting from the interaction of the flow over the blades and the blades' natural vibration tendencies. When flutter occurs, the unsteady aerodynamic forces on the blade, due to its vibration, add energy to the vibration, causing the vibration amplitude to increase. The vibration amplitude can become large enough to cause structural failure of the blade. The operable range, in terms of pressure rise and flow rate, of turbomachinery is restricted by various flutter phenomena.
Lower frequency vibration modes, i.e., the first bending mode and first torsion mode, are the vibration modes that are typically susceptible to flutter. In one approach to avoid or reduce flutter, it has been a conventional practice to increase the first bending and first torsion vibration frequencies of the blades, including utilizing mix-tuning principles that promote blade-to-blade differences in blade natural frequency and mode shape.
In highly loaded last row blades of typical power generation steam turbines, one strong contributor to aero-elastic instability is attributed to the shock associated with the supersonic expansion downstream of the blade passage throat, which may be referred to as shock induced flutter. Shock induced flutter may exist under either stalled or unstalled flow conditions, as is referenced to the presence or absence, respectively, of a gross separation of the flow about the airfoil surface as a result of inlet incidence angle effects. Under such conditions, the strength of the destabilizing forces associated with the shock flow field may be increased by the regularity of the blade-to-blade flow field behaviour.
The present invention provides an array of flow directing elements, such as blades, that include first and second flow directing elements or blades that operate to interrupt a regular element-to-element flow field, changing the flow field from a substantially symmetric flow field, formed when the flow directing elements are all the same, to a substantially asymmetric flow field created by forming the second flow directing elements with a dimensional characteristic that is different than a corresponding dimensional characteristic of the first flow directing elements. The terms “element-to-element flow field” and/or “blade-to-blade flow field”, as used herein, refers to a relationship, such as a flow field relationship, established between flow directing elements or blades located on a common row extending circumferentially around a rotor disk in a turbomachine.
In accordance with one aspect of the invention, an array of flow directing elements for use in a turbomachine is provided comprising a plurality of flow directing elements mounted to a rotor disk. Each of the flow directing elements includes a radially extending span dimension and a chord dimension extending substantially perpendicular to the span dimension. The plurality of flow directing elements comprise first flow directing elements forming a first set of flow directing elements and second flow directing elements forming a second set of flow directing elements. An element-to-element flow field defined between successive ones of the first set of flow directing elements is interrupted by the second set of flow directing elements to form an asymmetric element-to-element flow field around the array of flow directing elements.
In accordance with another aspect of the invention, an array of flow directing elements for use in a turbomachine is provided comprising a plurality of flow directing elements mounted to a rotor disk. Each of the flow directing elements includes a radially extending span dimension and a chord dimension extending substantially perpendicular to the span dimension. The plurality of flow directing elements comprises first flow directing elements forming a first set of flow directing elements and second flow directing elements forming a second set of flow directing elements. The second set of flow directing elements has a chord dimension defined by a value that is different than the value of a chord dimension measured at corresponding span-wise locations of the first set of flow directing elements.
In accordance with a further aspect of the invention, an array of flow directing elements for use in a turbomachine is provided to increase flutter stability, the array comprising a plurality of flow directing elements mounted to a rotor disk. Each of the flow directing elements includes a radially extending span dimension and a chord dimension extending substantially perpendicular to the span dimension. The plurality of flow directing elements comprises first flow directing elements forming a first set of flow directing elements and second flow directing elements forming a second set of flow directing elements. The second set of flow directing elements has a chord dimension defined by a value that is smaller than the value of a chord dimension measured at corresponding span-wise locations of the first set of flow directing elements to interrupt a shock field downstream of the flow directing elements and reduce shock induced flutter in the flow directing elements.
While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein:
In the following detailed description of the preferred embodiment, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific preferred embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.
Referring to the drawings, there is shown in
As shown in
In accordance with the illustrated embodiment, the center section of each blade 16 may also include a front standoff 34 and a rear standoff (not shown), where the front standoff 34 and rear standoff define mid-span snubber members, and where “front” and “rear” are referenced with respect to a turbine rotational direction. The mid-span snubber members each have a distal end defining respective snubber contact surfaces that form a small gap defining a snubber region therebetween.
In addition, a shroud portion 36 may be provided at the tip portion 32 of each of the blades 16. Each shroud portion 36 comprises a front end or contact surface 38 and an opposing rear end or contact surface 40. In the illustrated embodiment, the front and rear contact surfaces 38, 40 of adjacent blades 16 define an interlocking Z-shroud region comprising a small gap located between the contact surfaces 38, 40. When the turbine 10 is in use, the adjacent contact surfaces of the mid-span snubber members, and adjacent front and rear contact surfaces 38, 40 of adjacent shroud portions 32, may rub against each other as the blades 16 bend and twist during rotation of the rotor 14. As described herein, the blades 16 are shrouded blades that form a coupled blade structure; however, it should be understood that the present description may be considered substantially equally applicable to free standing blade structures, e.g., unshrouded blade structures.
As the steam 24 flows across the blades 16, from a leading edge 42 to a trailing edge 44, a flow field will be formed downstream of the trailing edge 44 that will have varying characteristics depending on the speed of the steam 24 passing through a given stage and the rotational speed of the blade 16. Further, the flow field may vary depending on the radial location on the blade 16, where locations along an inner span region of the blade 16 will tend to produce a subsonic flow field, and locations along an outer span region of the blade 16 will tend to produce a supersonic flow field. Flow fields comprising supersonic flows tend to produce aero-elastic instability that is evidenced by shock induced flutter of the blades 16.
Referring to
As seen with reference to
Flow fields having shock forces that create a flutter response in the blades 16a, 16b will generally occur within a range of exit Mach numbers, defined herein as a critical range of exit Mach numbers, such that the main parameter of concern with regard to the occurrence of flutter is the exit Mach number, which will generally determine the position at which the shock wave will impinge on the blades 16a, 16b. The shock waves defined within the critical range of exit Mach numbers comprises a range of positions generally defined between a first line 48, representing the shock wave produced by a lower limit exit Mach number, and a second line 50, representing the shock wave produced by an upper limit exit Mach number. The shock wave corresponding to the first line 48 will impinge on the blades 16a, 16b at axially forward locations 52a, 52b, respectively, and the shock wave corresponding to the second line 50 will impinge on the blades 16a, 16b at axially rearward locations 54a, 54b, respectively, where the locations 54b may generally correspond to the trailing edges 44b of the second blades 16b.
As seen in
Referring to
Since supersonic flow fields will generally occur at outer span portions of the airfoils 26a, 26b, the cut-back region 56 of the second airfoil 26b is defined starting at about 60% of the span length, where it blends with the profile of the unmodified first airfoil 26a, and continues to 100% of the span length, where it also blends with the profile of the unmodified first airfoil 26a. In the particular described embodiment, the trailing edge 44b may be cut back up to approximately 8%, e.g., by providing a generally corresponding reduction in the chord dimension C, at a radial location of about 70% to about 80% of the span length; and the trailing edge 44b may be cut back up to 4% at a radial location of about 90% of the span length.
The presently described blade array 20, providing alternating first and second blades 16a, 16b having normal and reduced chord dimensions C, respectively, operates to interrupt the flow field, changing the flow field from a substantially symmetric flow field, formed when the blades 16 are all the same, to a substantially asymmetric flow field. It should also be noted that the invention is not limited to the particular alternating arrangement of the blades 16a, 16b described herein and that the second blades 16b having modified chord dimensions may be provided in groups and/or may be separated by one or more of the first blades 16a having normal chord dimensions. Further, although a particular construction for the second airfoils 26b is described herein, the particular proportion(s) of the second airfoils 26b provided as cut-back areas 56 with a reduced chord dimension C may be varied to accommodate the particular operational conditions of the turbine.
The principles described herein may be particularly useful when implemented in a strongly coupled system, such as the above-described system including coupling components formed by adjacent contacting surfaces of the blades. Known techniques for reducing flutter by mix-tuning of blades, such as by tuning the natural frequency of blades, may be less effective in coupled systems as a result of the mechanical connection provided between the blades, and the presently described blade array may be provided to reduce the effect of shock forces that induce blade flutter. Further, the presently described blade array may be useful for reducing shock induced flutter in the blades of an uncoupled blade array, either in combination with other flutter and vibration reducing techniques, such as may be provided by altering the natural frequency of the blades, or when provided as a separate solution that may reduce the shock induced influence of adjacent blades in an array.
While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.
