Tunable gas turbine engine fan assembly
A gas turbine engine fan having blades embedded with a composite material. The size and the stiffness of the composite material control the vibratory characteristics of the blade without affecting the external blade profile. Accordingly, the gas turbine engine fan may use blades having similar external profiles with differing vibratory characteristics. The gas turbine engine fan can then be tuned using the blades with differing vibratory characteristics. Tuning the gas turbine engine fan minimizes the non-integral gas turbine engine fan vibrations.
The present invention relates to a gas turbine engine fan, and more particularly to tuning the gas turbine engine fan to reduce non-integral vibrations.
One goal of gas turbine engine design, specifically rotor and fan blade design, is to minimize non-integral vibrations caused by flutter and flow shedding. Tuning of the gas turbine engine fan has been shown to reduce non-integral vibrations in the gas turbine engine.
One known method of tuning the gas turbine engine fan involves varying the natural frequencies of individual fan blades by removing material from the blade edges. This method is commonly referred to as “clipping.” Removing material from the blade edges changes the natural frequency of the blade, and, in so doing, may reduce the non-integral vibrations in the fan. This method typically involves modifying the blade edges after the blade is manufactured.
Another method of tuning the gas turbine engine fan involves having fan blades with different thicknesses on the same rotor. Typically, this method uses fan blades on the rotor alternating between blades of different external profiles. Thus, both “thick” and “thin” blades are used and both “thick” and “thin” blades are manufactured.
Both of these approaches, “clipping” and “thick-thin,” have undesirable aerodynamic consequences. Preferably, fan blades on a rotor will have the same general profile.
Accordingly, there is a desire to tune a gas turbine engine fan without changing the fan blade profile. It is also desirable to incorporate the tuning method into the current manufacturing process of the gas turbine engine fan.
SUMMARY OF THE INVENTIONA turbine blade has a blade shell and a blade core each having a stiffness. Adjusting the stiffness of the blade shell or blade core may change the naturally frequency of the blade. In addition to the natural frequency of the blade, these adjustments may modify other tunable characteristics, e.g., weight.
The blade core is typically embedded in the blade shell. Accordingly, the blade core size can be modified without affecting the profile of the turbine blade. In addition, the blade core can be eliminated without changing the profile of the turbine blade. Thus, a turbine rotor disk containing multiple fan blades may contain fan blades with different sized blade cores. A turbine rotor disk may also contain one or more fan blades without blade cores.
A metal matrix composite (MMC) may be used as a blade core. An area of MMC material within the blade can be readily introduced to current turbine blade production techniques. A size of the MMC material may be adjusted to tune the gas turbine engine fan. Similarly, a stiffness of the MMC material may be adjusted to tune the gas turbine engine fan. Preferably, the MMC material is contained within alternating blades disposed upon the turbine rotor disk.
The present invention therefore provides a method of tuning a fan blade without altering the profile of the fan blade. In addition, the method of the present invention can be readily introduced to current blade manufacturing techniques.
BRIEF DESCRIPTION OF THE DRAWINGSThe various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the currently preferred embodiment. The drawings that accompany the detailed description can be briefly described as follows.
Referring to
Referring next to
As shown, the cavity 14 containing the blade core portion 18 is typically embedded within the blade shell portion 10. Embedding the cavity 14 enables blades 22 to be produced with similar profiles and dissimilar stiffnesses. For example, because the cavity 14 is embedded in the blade shell portion 10, blades 22 containing the cavity 14 may maintain the same profile as blades 22 not containing a cavity 14. In addition, the size of the cavity 14 may be changed without affecting the profile of the blade 22, provided the cavity 14 remains contained within the blade shell portion 10. Accordingly, a plurality of blades 22 may be disposed upon the rotor disk 30, and the blades 22 may or may not contain the blade core portion 18, as shown in
The blades 22 are ordinarily formed by diffusion bonding; two blade 22 halves are joined under extreme temperatures and pressures. A metal matrix composite (MMC) 26 material withstands these extremes. The MMC 26 is sandwiched between the two blade 22 halves as the blade 22 halves are bonded together.
Blades 22 may be made of Ti-6-4, while the MMC 26 may be Ti-6-4 with silicon carbide fiber additive. Varying the amount and orientation of the silicon carbide fiber additive alters the stiffness of the MMC 26. Of course, this invention extends to many other appropriate materials. In addition, stiffness can be varied by varying geometry, etc. Also while an alternate orientation of two different stiffnesses is disclosed, other arrangements of different stiffnesses could be utilized within the scope of this invention.
It should be understood that various alternatives to the embodiments of the invention described herein may be employed in practicing the invention. It is intended that the following claims define the scope of the invention and that the method and apparatus within the scope of these claims and their equivalents be covered thereby.
Claims
1. A rotor for a gas turbine engine, comprising:
- a rotor mounting a first group of blades;
- said rotor also mounting a second group of blades having a shell portion and a core portion; and
- wherein said second group of blades has a stiffness different from said first group of blades.
2. The rotor for a gas turbine engine of claim 1, wherein said core portion is embedded in said shell portion.
3. The rotor for a gas turbine engine of claim 1, wherein adjusting the stiffness of said core portion tunes said second group of blades.
4. The rotor for a gas turbine engine of claim 1, wherein adjusting said shell portion size and said core portion size tunes said second group of blades.
5. The rotor for a gas turbine engine of claim 1, further comprising a plurality of core portions, wherein said plurality of core portions has a plurality of stiffnesses.
6. The rotor for a gas turbine engine of claim 1, wherein said blade core portion comprises a composite material.
7. The rotor for a gas turbine engine of claim 1, wherein said composite material is a metal matrix composite.
8. The rotor for a gas turbine engine of claim 1, wherein said second group of blades are disposed between alternating ones of said first group of blades.
9. The rotor for a gas turbine engine of claim 1, wherein said rotor and said blades form part of a fan section for the gas turbine engine.
10. The rotor for a gas turbine engine of claim 1, wherein said first group of blades includes a shell portion and a core portion.
11. A gas turbine engine, comprising:
- a fan section;
- a compressor;
- a combustion section;
- a turbine;
- at least one of said fan and said turbine including a rotor having a plurality of blades, said plurality of blades comprising a material with a first stiffness and a material with a second stiffness within one or more of said plurality of blades; and
- wherein varying said second stiffness tunes said one or more of said plurality of blades.
12. The gas turbine engine of claim 11, wherein said material with a second stiffness is disposed within alternating ones of said plurality of blades.
13. The gas turbine engine of claim 11, wherein profiles of the plurality of blades are generally similar.
14. The gas turbine engine of claim 11, wherein said material with a second stiffness is a metal matrix composite.
15. The gas turbine engine of claim 11, wherein said material with a second stiffness is embedded in said material with a first stiffness.
16. The gas turbine engine of claim 11, wherein said material with a second stiffness comprises a composite material.
17. The gas turbine engine of claim 16, wherein said material with a second stiffness comprises a metal matrix composite.
18. The gas turbine engine of claim 11, wherein said rotor and said blades are part of said fan.
19. A method of tuning a gas turbine engine fan, comprising the steps of:
- a) mounting a first fan blade with a first stiffness to a rotor;
- b) mounting a second fan blade with a second stiffness to the rotor; and
- c) repeating steps a-b selectively mounting the first fan blade and the second fan blade creating a gas turbine engine fan with blades of selected stiffness.
20. The method of claim 19, further comprising the step of varying said first stiffness, wherein varying said first stiffness alters vibratory characteristics of said first fan blade.
Type: Application
Filed: Aug 9, 2005
Publication Date: Feb 15, 2007
Inventor: Robert Morris (Portland, CT)
Application Number: 11/200,359
International Classification: F01D 5/14 (20060101);