KR19990062719A - Rotor stage for gas turbine engine - Google Patents

Abstract

The rotor stage for a gas turbine engine includes a rotor disk and a plurality of rotor blades. The rotor blades extend radially outward from the outer radial surface of the disk and are separated from each other at some distance. This distance is optionally changed to increase the aerodynamic damping of the rotor blades.

Description

Rotor stage for gas turbine engine

FIELD OF THE INVENTION The present invention relates generally to gas turbine engine rotor assemblies and, more particularly, to apparatus for controlling vibration in rotor stages.

Most conventional rotor stages in a gas turbine engine include a plurality of rotor blades mechanically attached to a disk for rotating about an axis. The rotor blades typically have fir-tree or dovetail blade roots that engage in engagement slots disposed on the outer radial surface of the disk. A disadvantage of mechanically attached rotor blades is that significant stresses appear in the disc under load adjacent to the attachment slot. Increasing the disc outer diameter helps the distance between adjacent slots to minimize stress. Unfortunately, increasing the disk diameter increases the overall size and weight of the rotor stage. In recent years, relatively lightweight integrated bladed rotors (IBR) are more widely used. The blades of the IBR are formed integrally with the disk (including metallurgically attached blades) rather than mechanically attached to the disk. Integral blades are more effective at moving the load of the blade compared to conventional mechanical attachment designs. As a result, the size and weight of the rotor disk is advantageously small.

Conventional rotor stages are often adjusted to avoid vibrational response and damped to reduce any vibrational response that occurs. The turning is typically known to be adjusted where the natural frequency of the rotor stage changes to avoid the frequency of periodic forcing appearing in the operating environment of the rotor stage. Damping is generally known as an adjustment that is taken to minimize the vibrational response generated by periodic or aperiodic (optionally described) forcing action. The periodic force acts at a separate frequency and can produce a resonant response in the rotor blades when the frequency of the force matches the natural frequency of the rotor blades. On the other hand, aperiodic forcing does not occur at a particular frequency, but rather the rotor blades respond (deflect) in aperiodic form. In the absence of sufficient damping, both periodic and non-periodic stimulus forces can provide a high blade vibration response for all modes of vibration appearing in the operating speed range.

Mechanical, aerodynamic and material damping appears to be the three main types of damping potentially used in rotor stages. The material damping that occurs in conventional rotor stages and IBRs is the least effective of the three types and generally will not provide adequate damping of the rotor blades alone. On the other hand, mechanical damping has the greatest effect among three types and can be achieved by several different methods. In one method, the vibratory motion is damped by friction between the blade root and the disk slot, ie blade root damping. In another method, the friction device is attached to the rotor blades outward or inward so that motion is damped. In another example, blade-to-blade shrouds cause energy to disappear along the blade tip. However, this example of mechanical damping is not practical in having an optimal IBR due to the integral nature of the IBR blades. Neither the separate damper device between the IBR rotor blades and the disc or the device between adjacent IBR rotor blades is practical.

Aerodynamic damping is commonly known to turn into work between the rotor stage and the air passing through the rotor stage. For example, if the net work performed by air on the rotor blades exceeds that performed by the rotor blades in the air, the air adds energy to the blades. This results in an unstable condition, and blade vibrations can start and / or increase with magnitude and ultimate consequences for fatigue. On the other hand, when the net work performed by the rotor blades in air exceeds the work performed by the air on the rotor blades, the rotor blades dissipate energy into the airflow. This energy transfer away from the rotor blades represents a desirable state of aerodynamic damping.

Thus, what is needed can be used in IBRs, damping periodic coercive forces and aperiodic (optional) perturbations, and damping vibration responses in the rotor stage that are effectively damped in medium and low shaped rotor blades. Apparatus and / or method for

It is therefore an object of the present invention to provide a rotor stage for a gas turbine engine having a device for damping vibrations.

Another object of the present invention is to provide an apparatus for damping vibrations that can be used in an IBR.

Another object of the present invention is to provide an apparatus for effectively damping vibrations against vibrations generated by periodic forced action and aperiodic perturbation.

It is another object of the present invention to provide an apparatus for vibration damping that effectively damps vibrations in rotor blades of medium and low aspect ratios.

According to the invention, the rotor stage for a gas turbine engine comprises a rotor disk and a plurality of rotor blades. The rotor blades extend radially outward from the outer radial surface of the disk and are separated from each other at some distance. The distance is optionally changed to increase the airflow damping of the rotor blades.

The energy transferred from the observer's aerodynamic damping point to the rotor blades can be described as the unstable work performed by the blades in the air passing by the blades during the period of vibration, using Equation 1 below.

here, Denotes the difference in unstable air pressure acting on the suction and pressure side surfaces of the rotor blades at any point due to the effect of time as a result of the blade under oscillating motion, where W (x, y, z, t) Rotor blade deflection in any direction according to the The work expression is matched over time period T, where T is equal to the duration of one blade vibration. Positive work per cycle (indicated by the positive value of one equation) represents work performed on the blade by the passing air, i.e., work in an unstable state. Negative work per cycle (indicated by the negative value of one equation) represents the work in the desired state of aerodynamic damping, ie performed by the blades along with the passing air. The zero value of one equation is called a neutral condition, ie the blade will not accept or perform work.

Since the purpose of aerodynamic damping is to damp certain vibrational motions, it can be seen that the deflection function W (x, y, z, t) in Equation 1 can be considered not different. Thus, aerodynamic damping is an unstable pressure function Ensuring work can be achieved by adjusting the pressure, which is carried out on the blades and by the opposite blades. The unstable pressure difference acting on the rotor blades is a function of 1) the volume of air passing through the rotor blades and 2) the volume of air between adjacent rotor blades and 3) the relative motion of the adjacent blades. In the present invention, the volume of air between adjacent rotor blades is adjusted by selectively changing the distance between the rotor blades.

An advantage of the present invention is that a means for aerodynamic damping is provided. In any application, aerodynamic damping can be used to increase mechanical and / or material damping. In other applications where mechanical and / or material damping is limited (eg IBR), it may serve as the primary damping means of aerodynamic damping.

Another advantage of the present invention is that the rotor stage damping device is effective against vibrations generated by periodic forcing and aperiodic perturbation. The optional rotor blade spacing of the present invention allows the rotor blade to perform work by the air passing through it, regardless of whether the periodic forced action is subjected to aperiodic perturbation.

Another advantage of the present invention is that vibrations can be effectively damped in the rotor blades of medium and low aspect ratios. Medium and low aspect ratio rotor blades, which are typically attached or integrally formed, are particularly suitable for chordal modes of vibration. The optional rotor blade spacing of the present invention allows the rotor blades to damp the deflections caused by the complex current modes of periodic and aperiodic perturbation and vibration in principle.

Another advantage of the present invention is that no additional hardware, inner blade arrangement, or the like is required. The present invention also provides damping by selectively placing the rotor blades around the outer radial surface of the rotor disk. Those skilled in the art will appreciate that simplicity is typically the same as reliability.

These and other objects, features and advantages of the present invention will become apparent from the following detailed description of preferred embodiments of the invention, as shown in the accompanying drawings.

1 is a schematic cross-sectional view of a gas turbine engine,

2 is a schematic perspective view of a gas turbine rotor stage according to the prior art,

3 is a schematic perspective view of a gas turbine rotor stage according to the present invention;

Figure 4 shows a straight line representation of the rotor blades extending outward from the rotor disk according to the present invention, showing the spacing between the blades.

Explanation of symbols for main parts of the drawings

10 gas turbine engine 12 fan

18: burner 26: nozzle

34: disk 36: rotor blade

40: outer radial surface 42: distance between adjacent rotor blades

Referring to FIG. 1, the 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, a high pressure turbine 22, and an augmenter ( 24 and nozzles 26 arranged symmetrically about the axis of rotation 28. The fan 12 is in front of the nozzle 26 and the nozzle is at the rear of the fan 12. The fan 12 and the low pressure compressor 14 are connected to each other and driven by the low pressure turbine 20. The high pressure compressor 16 is driven by the high pressure turbine 22. The air actuated by the fan 12 will enter the low pressure compressor 14 as core gas or into the passage 30 outside the engine core as bypass air.

2 to 4, the rotor stage 32 includes a disk 34 and a plurality of rotor blades 36. The disc 34 includes a bore 38 (FIG. 3) concentric with the axis of rotation 28 and an outer radial surface 40. The rotor blades 36 extend radially outward from the outer radial surface 40 and may be attached to the disk 34 via conventional attachment methods (eg, fir or dovetail roots—not shown), or It can be integrally attached as part of an integrally bladed rotor (IBR).

In a typical rotor stage 32 (FIG. 2), the rotor blades 36 have a distance D (ie, equal to the circumferential length of the disc outer radial surface 40 divided by the number of rotor blades 36). D is equally spaced apart from each other by the distance between the centerlines of adjacent rotor blades). In the present invention (FIGS. 3 and 4), the distance 42 between adjacent rotor blades 36 (ie, the distance between rotor blades) is optionally modified to obtain an increase in aerodynamic damping. The amount of distance 42 between each altered rotor blade is not all but dependent on the application. In terms of the spacing distance D between uniform rotors, the distance 42 between adjacent rotor blades 36 is typically greater than 80% (0.80D) of the uniform spacing value and 120% of the uniform value D ( Less than 1.20D). However, the distance 42 between the rotor blades is preferably between 85% and 95% (0.85D-0.95D) of the uniform spacing value or between 115% and 105% (1.15D-1.05D) of the uniform spacing value D. Do. The optimum distance 42 (and thus the optimal damping) between adjacent rotor blades 36 is a function of the application environment and can be determined analytically or empirically. FIG. 4 schematically shows the spacing selectively altered by showing the distance between rotor blades equal to D, D- and D +, where D represents the distance between the uniform rotor blades, where D- is less than D Distance, and D + represents distance greater than D.

In some applications, most rotor blades 36 are evenly spaced apart from each other at a distance D, with only a few rotor blades 36 being non-uniformly spaced relative to adjacent rotor blades 36. In other applications, the distance 42 between most or all rotor blades will be uneven. However, for balance, it is desirable to achieve symmetry between the halves of the rotor disk 34 relative to the spaced rotor blades 36. For example, if the first blade is shifted a certain distance to the left of the evenly spaced position, the second blade is typically located 180 ° away from the first blade and from the uniformly spaced position by the same amount relative to the right side. It is desirable to move, ie, keep the rotor blades 180 ° offset from each other.

While the invention has been shown and described in accordance with the specific embodiments thereof, those skilled in the art will recognize that various changes may be made in the form and details of the invention without departing from the spirit and scope of the invention. For example, it may be desirable to have a distance between rotor blades less than 0.80D or a distance between rotor blades greater than 1.20D in any circumferential direction.

Thus, the present invention can be used in IBR and is effective against vibrations generated by periodic forced action and aperiodic perturbation, and effectively damps vibrations in the rotor braids of medium and low aspect ratios.

Claims (23)

A rotor stage for a gas turbine engine for rotating around a rotation axis, A rotor disk having a bore concentric with said axis of rotation and an outer radial surface having a circumferential length; And a plurality of rotor blades extending radially outward from the outer radial surface, wherein the adjacent rotor blades are separated from each other by the distance between the rotors, the distance between the rotors being selectively modified. The method of claim 1, And said selectively modified distance between rotors is at least 80% of the distance between uniform rotors. The method of claim 2, And wherein said selectively changed distance between rotors is at least 85% of the distance between said uniform rotors. The method of claim 3, wherein Wherein the distance between the selectively modified rotors is only 120% of the distance between the uniform rotors. The method of claim 4, wherein And said selectively modified distance between rotors is only 115% of the distance between said uniform rotors. The method of claim 2, Wherein the distance between the selectively modified rotors is only 120% of the distance between the uniform rotors. The method of claim 6, And said selectively modified distance between rotors is only 115% of the distance between said uniform rotors. The method of claim 6, And said rotor disk and rotor blades are integral blade type rotors. The method of claim 6, And each said rotor blade is offset 180 [deg.] From the other of said rotor blades. The method of claim 6, Wherein at least one of the distances between the rotors has been selectively modified. The method of claim 10, A rotor stage for a gas turbine engine, wherein the distance between a plurality of said rotors is optionally changed. The method of claim 11, Rotor stage for a gas turbine engine, wherein the distance between most of the rotors is optionally changed. The method of claim 2, And said rotor disk and rotor blades are integral blade type rotors. The method of claim 2, Each rotor blade is offset 180 ° from the other of the rotor blades. The method of claim 2, Wherein at least one of the distances between the rotors has been selectively modified. The method of claim 15, A rotor stage for a gas turbine engine, wherein the distance between a plurality of said rotors is optionally changed. The method of claim 1, And said rotor disk and rotor blades are integral blade type rotors. The method of claim 17, Each rotor blade is offset 180 ° from the other of the rotor blades. The method of claim 17, Wherein at least one of the distances between the rotors has been selectively modified. The method of claim 19, A rotor stage for a gas turbine engine, wherein the distance between a plurality of said rotors is optionally changed. The method of claim 1, Wherein the distance between the selectively modified rotors is only 120% of the distance between the uniform rotors. The method of claim 21, And said rotor disk and rotor blades are integral blade type rotors. The method of claim 21, And each said rotor blade is offset 180 [deg.] From the other of said rotor blades.