JPH11236804A - Rotor step for gas turbine engine - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a vibration damping device which can be used in a rotor integrally equipped with rotor blades, is effective for periodical push-in actions and aperiodical perturbation, and also is effective for vibration of rotor blade having middle or low grade aspect ratio. SOLUTION: Rotor steps 32 for a gas turbine engine include a rotor disc and a plurality of rotor blades 36. The rotor disc is provided with a hole centering on the axis of rotation and a surface 40 in radial direction on the outside. The rotor blades 36 are extended from the surface 40 in radial direction on the outside to the outside in the radial direction and are distributed around the surface 40. At least one rotor blade 36 is positioned aslant (α) selectively relative to other adjacent rotor blade 36. This slant rotor blade increases the aerodynamic damping of the rotor steps.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates generally to gas turbine engine rotor assemblies and, more particularly, to an apparatus for controlling rotor stage vibration.

[0002]

BACKGROUND OF THE INVENTION Many conventional rotor stages of a gas turbine engine include a plurality of buckets mechanically mounted on a disk for rotation about an axis. The blades typically have "Christmas tree" or "bee" shaped roots that fit into mating slots provided on the outer radial surface of the disk. A disadvantage of mechanically mounted blades is that significant stress develops in the loaded disk adjacent to the mounting slot. Increasing the outer diameter of the disk, and thus the distance between adjacent slots, helps to minimize stress.
However, increasing the disk diameter also increases the overall size and weight of the rotor stage.
For this reason, relatively lightweight “rotors with integral blades” have recently been widely used. The blades of the rotor with integral blades are integrally formed on the disk rather than being mechanically mounted on the disk (ie, the rotor includes the blades attached to the disk by metal bonding). . Such an integral moving blade is very effective in carrying the load of the moving blade as compared with a conventional mechanical mounting configuration. As a result, the size and weight of the rotor disk are advantageously minimized.

[0003] Conventional rotor stages are often rotated to eliminate the vibration response and damped to minimize the vibration response that occurs. Rotation generally refers to a procedure directed to changing the natural frequency of a rotor stage to remove the frequency of a periodic forcing function present in the operating environment of the rotor stage. Also, damping generally refers to a measure that minimizes the vibration response caused by a periodic or aperiodic (also referred to as random) pushing action. The periodic indentation acts at a discontinuous frequency, causing a vibration response in the blade when the frequency of the indentation reaches a match with the natural frequency of the blade. On the other hand, aperiodic indentation does not work at certain frequencies, but causes the blade to respond (distort) in an aperiodic manner. In the absence of sufficient damping, the two periodic and aperiodic excitation forces produce a high blade vibration response to all modes of vibration present in the operating speed range.

[0004] Mechanical, aerodynamic, and material damping represent three main types of potential damping used in rotor stages. Although material damping occurs in conventional rotor stages and also in rotors with integral blades, it is the least effective of the three types and generally provides sufficient damping for the blades by itself. It does not provide.
On the other hand, mechanical damping is the most powerful of the three types and can be achieved in several different ways. In one such method, the oscillating motion is damped, or "blade root," by friction between the blade root and the disk slot. In another approach, a friction device is mounted externally or internally to the blade to dampen oscillatory motion. In yet another approach, a shroud pairing adjacent blades is used to dissipate energy along the tip of the blade. However, the mechanical damping of these methods is useless for many rotor bladed rotors due to their rotor integrity. That is, the independent damping device between the rotor blade and the disk of the rotor with the integral rotor blade and the device between adjacent rotor blades of the integral rotor blade rotor are:
Both are useless.

[0005] Aerodynamic damping generally refers to the exchange of work between a rotor stage and air passing through the rotor stage.
If the net work provided by the air to the blades exceeds the net work provided by the blades to the air, for example, the air adds energy to the blades. This results in an unstable condition, in which the blades begin to oscillate, increasing in amplitude and eventually causing fatigue. On the other hand, if the net work provided by the blades to the air exceeds the net work provided by the air to the blades, for example, the blades will dissipate energy into the airflow. This transfer of energy away from the bucket results in a favorable state of aerodynamic damping.

From the foregoing, it can be seen that there is provided an apparatus and / or method for attenuating the vibration response of a rotor stage, which can be used for rotors with integral rotor blades, and which has a periodic pushing action and an aperiodic (random) effect. There is a need for an apparatus and / or method that attenuates critical perturbations and also effectively attenuates vibration of blades of medium and low aspect ratios.

[0007]

SUMMARY OF THE INVENTION Accordingly, it is an object of the present invention to provide a rotor stage for a gas turbine engine that includes a vibration damping device.

Another object of the present invention is to provide a vibration damping device that can be used for a rotor with an integral moving blade.

It is still another object of the present invention to provide a vibration damping device effective against vibration.

It is still another object of the present invention to provide a vibration damping device which is effective against vibration caused by a periodic pushing action and an aperiodic perturbation.

It is still another object of the present invention to provide a vibration damping device for effectively damping vibrations of blades having medium and low aspect ratios.

According to the present invention, there is provided a rotor stage for a gas turbine engine as described below. That is, the rotor stage for a gas turbine engine includes a rotor disk and a plurality of rotor blades. The rotor disk includes a hole centered on the axis of rotation and an outer radial surface. The blades extend radially outward from the outer radial surface and are distributed about the outer radial surface. And at least one blade is selectively beveled with respect to other adjacent blades. The one or more slanted blades increase the aerodynamic damping of the rotor stage.

[0013] In terms of aerodynamic damping, the energy transmitted to the blade can be determined by using the following equation (1) to provide the unstable energy imparted by the blade to the air passing along the blade during the vibration cycle. It can be said that it is work.

(Equation 1) put it here, Represents the unstable air pressure difference acting on the suction and pressure side surfaces of the blade at any point as a function of time, as a result of the blade undergoing an oscillating motion. Also, Represents the distortion of the bucket in any direction as a function of time. The work equation is integrated over the time period "T", where "T" is equal to the duration of one blade oscillation. The positive work per cycle (indicated by the positive value of the work formula) indicates the work given to the rotor blades by the passing air, that is, an unstable state. Negative work per cycle (indicated by the negative value of the work equation) indicates the work imparted to the passing air by the bucket, ie the favorable state of aerodynamic damping.
A work-type zero indicates the natural state, that is, the bucket is not given or given work.

Since the purpose of the aerodynamic damping is to damp the vibration of a predetermined mode, the distortion term of the above equation (1) is obtained. Can be considered immutable. Therefore, aerodynamic damping is not Can be achieved by assuring that work is provided by the bucket in opposition to the work provided to the bucket. The unstable pressure differential acting on the buckets is a function of 1) the air passing through the buckets, 2) the amount of air between neighboring buckets, and 3) the relative motion of the neighboring buckets. In the present invention, the unstable aerodynamic characteristics of air between adjacent blades occur by selectively skewing the chord lines of at least one blade with respect to the chord lines of other adjacent blades. It operates by increasing the aerodynamic damping that occurs.

[0015] An advantage of the present invention described above is that an aerodynamic damping device is provided. That is, in some applications, aerodynamic damping can be used to increase mechanical and / or material damping. Mechanical and / or
Or in other applications where material damping is limited (eg, rotors with integral blades), aerodynamic damping can be used as the primary damping device.

Another advantage of the present invention is that it provides a rotor stage damping device that is effective against periodic pushing and aperiodic perturbations. That is, the selective blade tilt according to the present invention allows the blade to provide work to the air passing through the blade, regardless of whether the blade is subjected to periodic pushing or aperiodic perturbations. Make it possible.

Yet another advantage of the present invention is that vibrations of blades of medium and low aspect ratios are effectively damped. That is, conventional mechanically attached or integrally formed medium and low aspect ratio blades are particularly susceptible to chord mode vibration. In contrast, the selective blade tilt according to the present invention allows the blade to attenuate distortions caused by major periodic and aperiodic perturbations in addition to full chord mode oscillations. .

Yet another advantage of the present invention is that the present invention does not require additional hardware, internal machining of blades or the like. That is, the present invention provides damping by selectively tilting at least one blade. One skilled in the art will recognize that simplicity is generally equal to trust.

The foregoing objects, features and advantages of the present invention will become more apparent from the following detailed description of preferred embodiments of the present invention with reference to the accompanying drawings.

[0020]

DETAILED DESCRIPTION OF THE INVENTION Referring to FIG. 1, a gas turbine engine 10 includes a fan 12, a low pressure compressor 14, a high pressure compressor 16, a combustor 18, a low pressure turbine 20, High pressure turbine 22 and augmentor 24
And a nozzle 26, which are provided symmetrically with respect to the axis of rotation 28. The fan 12 is a nozzle 26
And therefore the nozzle 26 is behind the fan 12. The fan 12 and the low-pressure compressor 14 are connected to each other, and are driven by a low-pressure turbine 22. High pressure compressor 1
6 is driven by a high-pressure turbine 22. Fan 12
The air, which is given a job by, enters the low pressure compressor 14 as "center gas" or enters the passage 30 outside the engine as "bypass air".

Referring to FIG. 2 to FIG.
Includes a disk 34 and a plurality of blades 36. Disc 34 includes a hole 38 (FIG. 2) centered on axis of rotation 28 and an outer radial surface 40.
The blades 36 extend radially outward from the outer radial surface 40 and are attached to the disk 34 by conventional mounting structures (eg, a Christmas tree or dovetail-not shown) or have integral blades. It can be provided integrally as a part of the rotor. Each bucket 36 has its leading edge 4
It has a chord line 42 extending between 4 and a trailing edge 46.

The conventional rotor stage 3 shown in FIGS. 2 and 3
2 are a plurality of moving blades 3 which are equally spaced apart from each other and distributed around the outer circumference of the rotor disk 34;
6. Each bucket 36 is in an array parallel to the other buckets 36 as evidenced by its parallel chord lines 42.
In contrast, in the present invention, one or more blades 36 are selectively beveled from a conventional parallel arrangement to achieve increased aerodynamic damping. Bucket 36
The amount by which is skewed depends on the particular application to be implemented.
In many applications, the blades 36 are angled from a conventional parallel arrangement by 5 degrees (5 °) in either direction (if the blades 36 are angled opposite in both directions). The total difference is 10 °). In the preferred embodiment, the buckets 36 are angled from the conventional parallel arrangement in no more than three degrees in either direction. FIG.
To illustrate the present invention, several buckets 36 are shown that are angled in both directions from a conventional parallel arrangement. The oblique angle is shown as "α" extending between the location of the chord line 42a associated with the parallel arrangement and the chord line 42b associated with the slanted bucket 36. The optimal diagonal (and hence optimal damping) of each bucket 36 of the rotor stage 32 is a function of the application and can be determined analytically or empirically.

In some applications, a majority of the blades 36 are maintained in a parallel arrangement, and only a small portion of the blades 36 are angled from the parallel arrangement. In other applications, most or all of the blades 36 are beveled from a parallel arrangement.

Although the present invention has been illustrated and described with reference to embodiments thereof, it will be understood by those skilled in the art that various changes can be made in the form and details without departing from the spirit and scope of the invention. There will be.

[Brief description of the drawings]

FIG. 1 is a schematic sectional view of a gas turbine engine.

FIG. 2 is a partial perspective view of a rotor stage of a conventional gas turbine engine.

FIG. 3 is a diagram showing a plurality of blades extending outward from a rotor disk linearly developed to show a conventional parallel blade arrangement, as evidenced by the parallel chord lines of the blades. is there.

FIG. 4 shows a plurality of blades extending outwardly from a rotor disk in a straight line to show a blade according to the invention oblique from a parallel arrangement, as evidenced by the oblique chord lines of the blades. FIG.

[Explanation of symbols]

 DESCRIPTION OF SYMBOLS 10 Gas turbine engine 12 Fan 14 Low pressure compressor 16 High pressure compressor 18 Combustor 20 High pressure turbine 22 Low pressure turbine 24 Augmentor 26 Nozzle 28 Rotation axis 30 Passage 32 Rotor stage 34 Disk 36 Blade 38 Hole 40 Outside radial surface 42, 42a , 42b chord line 44 leading edge 46 trailing edge α oblique angle

────────────────────────────────────────────────── ───

[Procedure amendment]

[Submission date] April 16, 1999

[Procedure amendment 1]

[Document name to be amended] Drawing

[Correction target item name] All figures

[Correction method] Change

[Correction contents]

FIG.

FIG. 2

FIG. 3

FIG. 4

 ──────────────────────────────────────────────────の Continuing on the front page (72) Inventor Bradford A. Cowles Hackbury Street 11654 Palm Beach Gardens, Florida 33654 USA

Claims (11)

[Claims] 1. A rotor stage for a gas turbine engine rotating about an axis of rotation, a rotor disk having a hole centered on the axis of rotation and an outer radial surface, and a radius from the outer radial surface. A plurality of blades extending outwardly in a direction and distributed about the outer radial surface and having chord lines, wherein at least one of the blades is selective with respect to other adjacent blades. A rotor stage for a gas turbine engine, wherein the rotor stage is beveled, thereby causing an increase in aerodynamic damping of the rotor stage. 2. A rotor stage for a gas turbine engine rotating about an axis of rotation, wherein the rotor disk has a hole centered on the axis of rotation and an outer radial surface, and a radius from the outer radial surface. A plurality of blades extending outwardly in a direction and distributed around the outer radial surface and having a chord line, wherein at least one of the blades has a chord line adjacent thereto. A rotor stage for a gas turbine engine, wherein the rotor stage is selectively beveled with respect to a chord line of the rotor, thereby causing an increase in aerodynamic damping of the rotor stage. 3. The rotor stage for a gas turbine engine according to claim 2, wherein a chord line of said at least one moving blade is slanted by no more than 10 ° from a chord line of said adjacent moving blade. Rotor stage. 4. The rotor stage for a gas turbine engine according to claim 3, wherein a chord line of said at least one moving blade is slanted by not more than 5 ° from a chord line of said adjacent other moving blade. Rotor stage. 5. The rotor stage according to claim 4, wherein said rotor stage is a rotor with integral moving blades. 6. The rotor stage for a gas turbine engine according to claim 4, wherein a chord line of said at least one moving blade is inclined by no more than 3 ° from a chord line of said adjacent other moving blade. Rotor stage. 7. The rotor stage according to claim 6, wherein said rotor stage is a rotor with integral moving blades. 8. The rotor stage for a gas turbine engine according to claim 2, wherein a chord line of said at least one moving blade is slanted by no more than 5 ° from a chord line of said adjacent other moving blade. Rotor stage. 9. The gas turbine engine rotor stage according to claim 8, wherein said rotor stage is a rotor with integral moving blades. 10. The rotor stage for a gas turbine engine according to claim 8, wherein a chord line of said at least one moving blade is inclined by no more than 3 ° from a chord line of said adjacent other moving blade. Rotor stage. 11. The rotor stage for a gas turbine engine according to claim 10, wherein said rotor stage is a rotor with integral moving blades.