Document
(12) United States Patent
(io) Patent No.: US 6,764,282 B2
Suciu et al. (45) Date of Patent: Jul. 20,2004
(54) BLADE FOR TURBINE ENGINE 3,661,475 A * 511972 Anderson et al. ....... 4161219 R 4,451,203 A 511984 Langley 4,460,315 A 711984 Tseng et al.
Inventors: Gabriel L. Suciu, Glastonbury, CT (75) 5,007,800 A 411991 Hacault et al.
(US); Michael Babu, Fairfield, CT 5,018,941 A 511991 Heurtel et al.
(US); James R. Murdock, Tolland, CT 5,022,822 A 611991 Sincere (US) 5,067,876 A 1111991 Moreman, I11 .......... 4161219 R 5,067,877 A * 1111991 Youssef .................. 4161220 R (73) Assignee: United Technologies Corporation, RE33,954 E 611992 Honda et al.
Hartford, CT (US) 5,131,814 A 711992 Przytulski et al.
5,310,318 A 511994 Lammas et al.
Subject to any disclaimer, the term of this ( * ) Notice: 5,370,501 A 1211994 Udal1 patent is extended or adjusted under 35 5,395,213 A 311995 Stenneler U.S.C. 154(b) by 0 days. 5,720,596 A * 211998 Pepperman ............. 4161220 R 5,860,787 A * 111999 Richards ................. 4161220 R 5,993,162 A * 1111999 Weisse et al. ........... 4161219 R (21) Appl. No.: 09/991,149 6,042,333 A 312000 Day 6,155,788 A 1212000 Beckford et al.
(22) Filed: Nov. 14, 2001 6,183,202 B1 * 212001 Ganshaw ................ 4161219 R (65) Prior Publication Data * cited by examiner US 200410062651 A1 Apr. 1, 2004 Primary E x a m i n e r a d w a r d K. Look (51) Int. C1.7 .............................. FOlD 5/32; FOlD 5/14 Assistant Examinerqichard A. Edgar (52) U.S. C1. .............................. 416/220 R; 4161223 A, (74) Attorney, Agent, or F i r m a r i a n J. Hamilla 4161243 (57) ABSTRACT (58) Field of Search ......................... 4161204 A, 219 R, 4161220 R, 223 A, 234, 243, 248 A blade for a turbine engine having a centerline. The blade comprises: a root section extending at an angle relative to the (56) References Cited centerline; and an airfoil section extending from the root section. The root section is directly adjacent said airfoil U.S. PATENT DOCUMENTS section. In other words, the blade is neckless. The blade is part of a rotor assembly, and is preferably a fan blade.
2,769,611 A * 1111956 Schwarzkopf .......... 4161219 R 2,775,426 A * 1211956 Barrett, Jr. et al. ..... 4161219 R 3,323,710 A 611967 Daly 23 Claims, 10 Drawing Sheets
U S . Patent Jul. 20,2004 Sheet 1 of 10 US 6,764,282 B2
U S . Patent Jul. 20,2004 Sheet 2 of 10 US 6,764,282 B2
FIG.2a
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F K 2 b
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100"
f
131"
FIG.3
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I U
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FIG.7
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FIG. 11 a
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U S . Patent Jul. 20,2004 Sheet 9 of 10 US 6,764,282 B2
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FIG. 14
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U S . Patent Jul. 20,2004 Sheet 10 of 10 US 6,764,282 B2
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FIG.17 539
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It is a further object of the present invention to reduce the BLADE FOR TURBINE ENGINE size of the disk.
STATEMENT OF GOVERNMENTAL RIGHTS It is a further object of the present invention to increase mass flow through the fan while keeping engine diameter The U.S. Government may have rights in this invention constant.
S pursuant to NASA contract NAS3-98005.
It is a further object of the present invention to reduce TECHNICAL FIELD blade length while keeping mass flow through the fan constant.
This invention relates to a blade for a turbine engine.
It is a further object of the present invention to decrease Specifically, the invention relates to a fan blade for a gas the kinetic energy of the blade during a blade loss event. 10 turbofan engine.
It is a further object of the present invention to reduce the BACKGROUND OF THE INVENTION size of the containment structure used to confine a released blade.
FIG. 1 provides a cross-sectional view of a gas turbofan It is a further object of the present invention to decrease engine 50 in a nacelle N. Briefly, air enters an inlet 51 in the 1s the unbalanced load on the rotor after a blade loss event.
nacelle N. A fan section 53 compresses the air entering the It is a further object of the present invention to reduce the inlet 51. The fan section 53 also splits the air into a primary, structural requirements of the engine and aircraft, such as the or core, engine flow C and a secondary, or bypass, flow B.
size of the engine cases, struts, flanges, supports, mounts and From this point, these flows will travel different paths engine pylons. through the engine.
These and other objects of the present invention are The core engine flow C enters a compressor section of the achieved in one aspect by a blade for a turbine engine having engine. Typically, the compressor section includes a low a centerline. The blade comprises: a root section extending pressure compressor 55 and a high pressure compressor 57.
at an angle relative to the centerline; and an airfoil section The compressor section increases the pressure of the air to extending from the root section. The root section is directly aid in the combustion cycle.
adjacent said airfoil section The compressed core engine flow C then enters a diffuser/ 2s These and other objects of the present invention are combustor section 59. The diffuser decreases the velocity of achieved in another aspect by a blade for a turbine engine the core engine flow C and further increases pressure. The having an axial direction. The blade comprises: an axially combustor section 59 mixes the core engine flow C with fuel oriented root section; and an airfoil section extending from (not shown) and combusts the mixture.
the root section. The blade does not have a neck between the The gases from the combustor section 59 then enter a root section and the airfoil section.
turbine section. Typically, the turbine section includes a high These and other objects of the present invention are pressure turbine 61 connected to the high pressure compres- sor 57 and a low pressure turbine 63 connected to the low achieved in another aspect by a rotor assembly for a turbine engine having an axial direction. The rotor assembly pressure compressor 55 and fan.
includes: a disk having a plurality of axially oriented After driving the high pressure turbine 61 and the low 3s grooves; and a plurality of neckless blades. Each blade has pressure turbine 63, the core engine flow C exits the engine a root section with a continuous enlarged head for placement 50 through a nozzle 65. The core engine flow C through the within a corresponding groove.
nozzle 65 produces thrust.
These and other objects of the present invention are The bypass flow B avoids the core engine. Instead, the achieved in another aspect by a turbofan engine having a bypass flow B travels around the core engine by following flow path. The engine comprises: a fan section; a compressor the fan section 53 and exiting through a nozzle 67. The section; a burner section; a turbine section; and an exhaust bypass flow B through the nozzle 67 also produces thrust.
section. The fan section includes a disk and a plurality of The thrust produced by the bypass flow B in high bypass blades secured thereto. The outer surface of the disk and a ratio turbofans can account for a significant portion (e.g. 75 portion of the blades define an inner boundary of the flow percent) of total engine thrust.
4s path.
As thrust requirements increase, designers typically increase the diameter of the engine 50. While producing BRIEF DESCRIPTION OF THE DRAWINGS greater thrust, the larger engine adds weight to the aircraft.
Other uses and advantages of the present invention will A portion of the weight increase occurs directly within the become apparent to those skilled in the art upon reference to engine. For example, the larger engine has larger and heavier the specification and the drawings, in which: so fan blades that require, for example, heavier disks, bearings FIG. 1 is a cross-sectional view of a gas turbofan engine; and supports. A portion of the weight increase also occurs FIG. 2a is a perspective view of a conventional rotor indirectly. For instance, larger fan blades require a stronger assembly; containment structure to absorb a blade loss. Also, a larger FIG. 2b is a perspective view of another conventional engine requires a stronger pylon on the aircraft and larger rotor assembly; 5s struts, flanges, supports and mounts on the nacelle.
FIG. 2c is a perspective view of another conventional Thus, a need exists for keeping weight increases to a rotor assembly; minimum. In fact, a preference exists for reducing weight FIG. 3 is a schematic showing the arrangement of various whenever possible.
components of a turbine engine using a conventional disk DISCLOSURE OF THE INVENTION and blade such as those shown in FIGS. 2a-c; 60 FIG. 4 is a front view of one alternative embodiment of It is an object of the present invention to reduce engine a rotor assembly of the present invention; weight.
FIG. 5 is a side view of the rotor assembly of FIG. 4; It is a further object of the present invention to reduce fan FIG. 6 is a perspective view of a portion of the rotor blade weight.
assembly of FIG. 4; 65 It is a further object of the present invention to reduce the FIG. 7 is a perspective view of a portion of the disk used size of the retention structure that secures the fan blade to the in the rotor assembly of FIG. 4; disk.
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FIG. 8 is a perspective view of a portion of the blade used Similar to the dove tail 115 of FIG. 2a, the dove tail 115" in the rotor assembly of FIG. 4; extends generally parallel to the axial centerline A of the engine.
FIG. 9 is a cross-section of a portion of a turbine engine incorporating the rotor assembly of FIG. 4; FIG. 3 schematically displays the common features of the s conventional rotor assemblies shown in FIGS. 2a-c. In these FIG. 10 is a perspective view of a portion of a disk used rotor assemblies, the dovetails 115, 115', 115" extend along in an alternative embodiment of the rotor assembly; a line D generally parallel to the axial direction. Although FIG. l l a is a front perspective view of a portion of a blade generally parallel to the axial direction, the dovetails 115 can used in the alternative embodiment of the rotor assembly; have a slight radial component. An angle a shows the radial FIG. l l b is a rear perspective view of the blade used in the deviation of line D from axial centerline A. Angle a can be alternative embodiment of the rotor assembly; range approximately between 0" and 10".
FIG. 12 is a cross-section of a portion of a turbine engine Also in the aforementioned rotors, the inner boundary of incorporating the alternative embodiment of the rotor assem- the core engine flow path extends along a line I. Although bly; generally extending in the axial direction, line I also has FIG. 13 is a cross-section of a portion of a turbine engine 15 radial component. An angle fl shows the radial deviation of incorporating another alternative embodiment of the rotor line I from axial centerline A. Angle fl can range approxi- assembly; mately between approximately 10" and 25".
FIG. 14 is a perspective view of another alternative As seen from FIG. 3, the difference between the angles of embodiment of the rotor assembly; lines I and D can reach approximately 25". Atransition area, FIG. 15 is a plan view of a portion of a disk used in the or neck, typically occupies the area between the dovetail and 20 .
rotor assembly of FIG. 14; inner boundary of the core engine flow path. The neck, FIG. 16 is a perspective view of a portion of a blade used however, is not a "working" part of the blade. The neck in the rotor assembly of FIG. 14; and neither helps retain the blade in the disk nor compresses the core engine flow C. Such a non-working part only adds FIG. 17 is a plan view of a portion of the rotor assembly weight to the engine.
of FIG. 14.
The present invention does not use a transition area or BEST MODE FOR CARRYING OUT THE neck. The remaining figures describe various alternative INVENTION embodiments of the present invention.
FIGS. 2a-c display various conventional rotor assem- FIGS. 4-6 display several views of one alternative 2a, rotor assembly 100 includes a disk 101 blies. In FIG. embodiment of a rotor assembly. The rotor assembly 200 with a plurality of grooves 103. The grooves 103 extend 30 includes a disk 201 and blades 203. FIG. 7 provides a generally parallel to an axial centerline A of the turbine perspective view of a portion of the disk 201.
engine. Each groove 103 receives a corresponding blade The disk 201 has an annular shape, with a front face 205, 111.
a rear face 207 and an outer surface 209. As clearly seen in The distal, or outer, region of the blade 111 includes an FIG. 5 , the outer diameter of the disk 201 increases from the airfoil 113. The proximal, or inner, region of the blade 111 35 front face 205 to the rear face 207. Although the figures (see, includes a root section having a dove tail 115. In order to e.g., FIG. 9) show the outer surface 209 of the disk 201 engage a corresponding groove 103, the dove tail 115 following a curvilinear path between the front face 205 and extends generally parallel to the axial centerline A of the the rear face 207, the outer surface could follow any suitable engine. The dove tail 115 also includes a notch 117. path, such as rectilinear.
Between the dove tail 115 and the airfoil 113, a platform The outer surface 209 substantially defines the inner 40 119 extends from the blade 111. The platform 119 extends in boundary of core engine flow path. The outer surface 209 a radial direction, at an angle to the axial centerline A of the has a plurality of grooves 211 generally extending between engine. The platform 119 abuts the platforms on adjacent the front face 205 and the rear face 207. A shoulder 213 blades. The platforms 119 create the inner boundary for the exists adjacent the grooves 211 along the front face 205. The core engine flow path. A neck 121 resides between the dove 45 grooves 211 define one half of the retention structure that tail 115 and the platform 119. secures the blades 203 to the disk 201.
A split lock ring 131 helps secure the blades 111 to the As best seen in FIG. 9, the grooves 211 extend at an angle disk 101. After placing the blades 111 in the grooves 103, the to the axial centerline A. The grooves 211 preferably travel split lock ring 131 is placed within the notches 117 of the a rectilinear path from the front face 205 to the rear face 207.
dove tails 115. The grooves 211 generally follow the outer surface 209 of the disk 201. Since the outer surface 209 may follow a FIG. 2b displays a similar rotor assembly 100' having a 211 may have local- disk 101' and blade 111'. Rather than using the notch 117 and curvilinear path, the rectilinear grooves ized areas that are slightly non-parallel (e.g. up to approxi- split lock ring 131 retention arrangement of FIG. 2a, the mately 5").
dove tail 115' includes a tab 133' that abuts the face of the disk 101'. When a cone segment (not shown) secures to the As seen in FIG. 7, the grooves 211 also have an arcuate forward face of the rotor loo', the tab 133' becomes wedged 55 shape. This accommodates the complex geometry of an between the rotor 100' and the cone segment. airfoil section 215 of the blade. FIG. 8 displays a portion of the blade 203.
115 of FIG. 2a, the dove tail 115' Similar to the dove tail extends generally parallel to the axial centerline A of the The airfoil 215 resides at the distal, or outer, region of the engine. blade 203. The blade 203 also includes a root section having 6o a dovetail 217 for insertion into a corresponding one of the FIG. 2c displays another rotor assembly 100". Similar to grooves 211. The root section of the blade 203 also includes the rotor 100, rotor assembly 100" uses a split lock ring 131" a tab 219.
to help retain the blades 111" to the disk 101". Differently than the blades in FIGS. 2a and 2b, the blade 111" does not The dovetail 217 defines the other half of the retention use a platform to define the inner boundary for the core structure used to secure the blades 203 to the disk 201.
engine flow path. Rather, inserts 135" are placed between 65 Although the figures show the retention structure as the adjacent blades 111". The inserts 135" define the inner groove 211Iidove tail 217, any other arrangement suitable to boundary for the core engine flow path. secure the blade 203 to the disk 201 could be used. For
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example, the retention structure could use a fir tree arrange- The rotor assembly uses a different retention feature than ment rather than the dovetail 217. rotor assembly 200 at the rear end of the rotor assembly.
Specifically, the engine uses fasteners F to secure the disk
221 between the root ne blade 203 includes a
301 to an flange 331 On the low pressure section and the airfoil 215. After installing the blades 203 the rear Of the 331 extends flange into the grooves 211 of the disk 201, the platform 221 serves s 325. The to fill in the gaps. In other words, the platforms 221 define s ~ ~ ~ e ~ ~ ~ ; ,& , " b i t : : a small portion of the inner boundary of the core engine flow disk 301. To prevent core engine flow disturbances, the path. As Seen in the platforms 221 are flush with the apertures are subsequently filled with a suitable sealant outer surface 209 of the disk 201.
material.
The platforms 221 of the present invention are narrower i o FIG. 13 displays another alternative embodiment of the than COnventionalblades. Generally speaking, the Platforms rotor assembly. Since the rotor assembly is similar to the 221 are narrower because the outer surface 209 of the disk aforementioned rotor assemblies, only the differences will 201 defines the majority of the inner boundary of the core be discussed. Similar features will use the same reference engine flow path. The blade 203 is considered neckless character, except for a change in the hundreds digit.
because the root section of the blade 203 transitions directly The main difference between this rotor assembly and the into the airfoil section. other embodiments resides in the rear retention feature. As FIG. 9 displays the rotor assembly 200 fully installed in Seen in FIG. 13, the rotor assembly uses a second split lock an engine. To reach this point, assembly proceeds as follows. ring 433. T O aCCOmmhte the lock ring 433, the rear ofthe within the grooves 211 of disk 401 includes a plurality of extensions 435 similar to The blades 203 are serially the disk 201 until the tabs 219 abut the shoulder 213. Then, those on the front of the disk 401 used to retain lock ring a split lock ring 223 is in a gap formed between the 2o 423. The extensions 435 flank the grooves 411 and provide blades 203 and the disk 201 to prevent forward movement a gap between the rear ofthe disk 401 and the blade 403. The of the blades 203. Differently than conventional blades, the lock ring 433 is Placed in the gap to Prevent rearward tabs 219 of the present invention prevent rearward move- m ~ e m e n t of the blades 401.
ment of the blades created by rotation. Rearward movement FIGS. 14-17 display another alternative embodiment of occurs because the grooves 211 extend in both an axial 25 the rotor assembly. Since the rotor assembly is similar to the direction and a radial direction. other rotor assemblies, only the differences will be dis- ne rotor assembly 100 is then secured to a low pressure cussed. Similar features will use the same reference compressor 225 and a cone segment 227 using conventional character, except for a change in the hundreds digit.
The main difference between this rotor assembly and the techniques.
benefits, 30 other embodiments resides in the retention features. As seen
ne use of the present invention has
an 152 the groove 511 in the disk 501 The primary benefit of using the present invention is reduced in 505. The fan blade weight. This primary benefit produces numerous enlarged section 537 adjacent the front face enlarged section 537 receives an extension 539 on the other benefits.
503. The platform 521 and the First, a lighter fan blade can use smaller retention struc- platform 521 Of the ture to retain the blade to the disk, A smaller retention 35 extension 539 create a wedge shape at the front Of the l6 the extension 539 On the platform 521.
structure enables the use of a smaller diameter disk, A l4 and l7 show the 503 secured to the disk smaller diameter disk allows increased mass flow through 501. The wedge-shaPed section Of the 503 abuts the the fan (assuming constant engine diameter). Alternately, a CorresPondinglY shaped enlarged section 537 of the groove smaller diameter disk allows decreased engine diameter 511 in the disk 501. The wedge shape prevents rearward while providing the same mass flow through the fan.
movement of the blade 503. As with the other embodiments, Second, a lighter fan blade decreases the kinetic energy of a lock ring (not shown) placed between the disk 501 the blade during a blade loss event, The lower energy and the 503 prevents forward movement Of the produces enables the reduction in size of the containment 503.
structure used to confine the released blade.
The present invention has been described in connection Third, a lighter fan blade decreases the unbalanced load 45 with the preferred embodiments Of the various figures. It is on the rotor after a blade loss event, A smaller unbalanced to be understood that Other embodiments may be load reduces structural requirements, such as the size of the Or modifications and additions may be made to the engine cases, struts, flanges, supports, mounts and engine described embodiment for performing the same function of pylons.
the present invention without deviating therefrom.
display another alternative embodiment of FIGS, 50 Therefore, the present invention should not be limited to any the rotor assembly. Since the rotor assembly is similar to sing1e embodiment? but rather construed in breadth and rotor assembly 200, only the differences will be discussed, scope in accordance with the recitation of the appended Similar features will use the same reference character, claims.
except for a change in the hundreds digit.
What is claimed is: lo a disk 301' The disk 301 has the Same 1, Ablade for a turbine engine having a centerline and a features as disk 201. The disk 301, however, includes one 55 flow path, the blade comprising: additional feature. The disk 301 includes apertures 329 a root section extending at an to the extending between rear face 307 and outer surface 309. The centerline; apertures are located between adjacent grooves 311.
an airfoil section extending said root section; and FIG, 11 displays a blade 303, The blade 303 has the Same a platform between said root on and said airfoil section; features as blade 203, except for the tab 219 on the dovetail 60 217. Blade 303 merely has a dovetail 319. wherein said root section and aid platform follow the flow path.
FIG. 12 displays the rotor assembly fully installed in an engine. As was described earlier with the rotor assembly 2. The blade as recited in claim 1, wherein the blade is a 200, the dovetails 319 of the blades 303 are inserted into the fan blade.
grooves 311 of the disk 301. Then, the split lock ring 323 is 65 3. The blade as recited in claim 1, further comprising a placed at the front of the rotor assembly in the gap between means for preventing rearward movement of the blade when the disk 301 and the blade 303. mounted to a disk.
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4. The blade as recited in claim 1, wherein said platform a plurality of neckless blades, each having a root section has a wedge shape to prevent rearward movement of the with a continuous enlarged head for placement within blade. a corresponding on of said grooves and a platform; 5. The blade as recited in claim 1, wherein said root wherein said root section follows the flow path and said platform defines a boundary of the flow path.
section is continuous. 5 6. The blade as recited in claim 5 , wherein said root 16. The rotor assembly as recited in claim 15, wherein section comprises an enlarged head. said blades are fan blades.
7. The blade as recited in claim 6, wherein said enlarged 17. The rotor assembly as recited in claim 15, wherein said enlarged head has a dovetail or fir tree shape.
head has a dovetail or fir-tree shape.
18. The rotor assembly as recited in claim 15, further 8. A blade for a turbine engine having an axial direction comprising a means for preventing rearward movement of and a flow path, the blade comprising: said blade when mounted to said disk.
an axially oriented root section following the flow path, 19. The rotor assembly of claim 15, further comprising a wherein, at said blade, the flow path extends at an angle lock ring upstream of said blades and a lock ring down- to the axial direction; stream of said blades to secure said blades to said disk.
an airfoil section extending from said root section; and 20. The rotor assembly of claim 15, wherein said disk 1s includes apertures therein, within the flow path, for securing a platform between said root section and said airfoil said disk to a flange of a downstream component.
section, and following the flow path; 21. A turbofan engine having a centerline and a flow path, wherein said blade does not have a neck between said root comprising: section and said airfoil section.
a fan section having a disk ant a plurality of blades 9. The blade as recited in claim 8, wherein the blade is a 2o secured thereto; fan blade.
a compressor section; 10. The blade as recited in claim 8, further comprising a a burner section; means for preventing rearward movement of the blade when a turbine section; and mounted to a disk.
an exhaust section; 11. The blade of claim 8, wherein said platform has a zs wedge shape to prevent rearward movement of the blade. wherein an outer surface of std disk and platforms on said 12. The blade as recited in claim 8, wherein said root blades define an inner boundary of the flow path, the section is continuous.
flow path extending at an angle relative to the center- 13. The blade as recited in claim 12, wherein said root line.
section comprises an enlarged head.
30 22. The engine as recited in claim 21, wherein said 14. The blade as recited in claim 13, wherein said platforms are flush with said outer surface of said disk.
enlarged head has a dovetail or fir tree shape.
23. The engine as recited in claim 21, wherein said blades 15. A rotor assembly for a turbine engine having an axial each have a retention section received within a correspond- direction and a flow path, comprising: ing groove in said disk, said retention section extending a disk having a plurality of axially oriented grooves that 3s parallel to said inner boundary of said flow path.
follow the flow path, wherein, at said disk, the flow
* * * * *
path extends at an angle to the axial direction; and
UNITED STATES PATENT AND TRADEMARK OFFICE
CERTIFICATE OF CORRECTION
PATENT NO. : 6,764,282 B2 Page 1 of 1 DATED : July 20,2004 INVENTOR(S) : Gabriel L. Suciu et al.
It is certified that error appears in the above-identified patent and that said Letters Patent is hereby corrected as shown below: Column 6,
Line 59, after “extending” insert -- from --
Line 60, delete “on” and insert -- section --
Line 61, delete “aid” and insert -- said --
Column 8,
Line 3, delete “on” and insert -- one --
Signed and Sealed this
Fifth Day of April, 2005
JON W. DUDAS Director of the United States Patent and Trademark Ofice