Appeal No. 2002-2148 Page 5
Application No. 09/627,143
Mannava's invention relates to gas turbine engine rotor blades and, more
particularly, to blade tips having localized compressive residual stresses imparted by
laser shock peening. Mannava teaches (column 1, line 59-66) that
[t]he burrs, nicks, and tears, hereinafter referred to as the tip damage, become
the source of high stress concentrations or stress risers and may severely limit
the life of the blades due to High Cycle Fatigue (HCF) from vibratory stresses
discussed above. It is also expensive to refurbish and/or replace compressor and
turbine blades and, therefore, any means to enhance the rotor capability and, in
particular, to extend aircraft engine blade life is very desirable.
Mannava then states (column 2, lines 40-49) that
[t]he present invention provides a gas turbine engine blade having at least
one laser shock peened surface along the tip of the blade and a region of deep
compressive residual stresses imparted by laser shock peening (LSP) extending
from the laser shock peened surface into the blade. Preferably, the blade has
laser shock peened surfaces on both suction and pressure sides of the blade
wherein both sides were simultaneously laser shock peened. The present
invention can be used for new, used, and repaired compressor and turbine
blades.
Figure 2 of Mannava illustrate a compressor blade having an airfoil 34 extending
radially outward from a blade platform 36 to a blade tip 38. The compressor blade
includes a root section 40 extending radially inward from the platform 36 to a radially
inward end 37 of the root section 40. At the radially inward end 37 of the root section
40 is a blade root 42 which is connected to the platform 36 by a blade shank 44. A
chord C of the airfoil 34 is the line between the leading edge LE and trailing edge TE at
each cross section of the blade as illustrated in Figure 4. The airfoil 34 extends in the
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