NBSIR 73-284

Examination of United Aircraft of

Canada Compressor Turbine Disk and

Blade Assembly

T. Robert Shives

Mechanical Properties Section

Metallurgy Division

Institute for Materials Research

National Bureau of Standards

Washington, D. C. 20234

October 3, 1973

Failure Analysis Report

Prepared for Bureau of Aviation Safety

National Transportation Safety Board

Department of Transportation

Washington, D. C. 20591

NBSIR 73-284

EXAMINATION OF UNITED AIRCRAFT OF

CANADA COMPRESSOR TURBINE DISK AND

BLADE ASSEMBLY

T. Robert Shives

Mechanical Properties Section

Metallurgy Division

Institute for Materials Research

National Bureau of Standards

Washington, D. C. 20234

October 3, 1973

Failure Analysis Report

"This document has been prepared for the use of the Bureau of Aviation Safety, National Transportation

Safety Board, Department of Transportation,

Washington, D. C. Responsibility for its further use

rests with that agency. NBS requests that if release

to the public is contemplated, such action be taken

only after consultation with the Office of Public

Affairs at the National Bureau of Standards."

Prepared for Bureau of Aviation Safety

National Transportation Safety Board

Department of Transportation

Washington, D. C. 20591

or

U. S. DEPARTMENT OF COMMERCE, Frederick B. Dent, Secretary

WATIOIMAL BUREAU OF STANDARDS, Richard W. Roberts, Director

TABLE OF CONTENTS

Page

SUMMARY

.

iii

1. GENERAL INFORMATION 1

1.1 Reference 1

1.2 Accident Information 1

1.3 Part Submitted 1

2. PURPOSE 1

3. RESULTS OF EXAMINATIONS AND TESTS 1

3.1 Visual and Macroscopic Examination 1

3.2 Fractographic Examination 3

3.3 Hardness Measurements 4

3.4 Metallographic Examination 4

4. DISCUSSION OF RESULTS 5

5. CONCLUSIONS 6

6. ACKNOWLEDGEMENT 6

FIGURES

1. Compressor turbine disk and blade assembly as

received 7

2. Part of the side of the turbine disk and blade

assembly containing the spline 8

3. Part of the compressor turbine disk and blade

assembly showing blade base askewness 9

4. Part of the compressor turbine disk hub 10

5. One of the fir trees of the disk and two blade

fir trees 11

6. Fracture surfaces of the three fractured disk

fir trees 12

i

Page

FIGURES (Continued)

7. Scanning electron photomicrograph showing

intergranular cracking 13

8 . Scanning electron photomicrograph showing

intergranular cracking 14

9. Scanning electron photomicrograph showing

dimpled rupture 15

10. Circumferential section through assembly 16

11. Crack regions A and B from figure 10 at

higher magnification 17

12. Etched circumferential section in crack

region B of one of the disk fir trees 18

13. Etched circumferential section showing

depleted zone 19

ii

SUJ^WARY

A damaged compressor turbine disk and blade assembly

from an aircraft engine was examined primarily to determine

the mode of fracture of three completely fractured disk fir

trees. Cracks had initiated and propagated for some

distance in an intergranular mode with the final fracture

being ductile overload. In addition to the three fractured

disk fir trees, the assembly had suffered considerable

mechanical damage. This damage was likely caused by the

assembly contacting other components of the system or a foreign object. The microstructure of the disk material

appeared normal for Waspaloy in the specified heat treated

condition and the hardness of the disk met specifications.

A depleted zone was found at the surfaces of the disk fir

trees and along the surfaces of some of the cracks in the

disk fir trees. The assembly thus had apparently been

subjected to the service atmosphere for a time after some

of the cracks had formed.

iii

Examination of United Aircraft of Canada

Compressor Turbine Disk and Blade Assembly

1 . GENERAL INFORMATION

1 . 1 Reference

Bureau of Aviation Safety, National Transportation

Safety Board, Department of Transportation, Washington, D. C.

Request by Mr. JVichael L. Marx, dated June 8, 1973. The

Bureau of Aviation Safety had been requested by the National

Board of Aviation, Helsinki, Finland, to provide technical

assistance in the examination of the assembly.

1 . 2 Accident Information

KAR-AIR OY DeHavilland Twin Otter DHC-6, series 200, serial

number 66. The accident occurred between Oulu and Kuusamo,

Finland, on February 5, 1973.

1 . 3 Part Submitted

One failed aircraft engine compressor turbine disk and

blade assembly. United Aircraft of Canada, Ltd., part number

3013511, was submitted for examination. Three fir tree regions

of the disk had fractured and separated from the assembly.

The separated parts of the disk and the blades which had been

attached to them were not submitted, except for a small piece

of one blade which remained attached to the assembly (arrov/ A,

figure 1)

.

2 . PURPOSE

The Bureau of Aviation Safety requested the NBS Mechanical

Properties Section to document the damage to the disk, perform

a f ractographic examination of the three fractured fir tree

areas of the disk, perform a metallographic examination of the

disk material, and make hardness measurements on the disk

material

.

3 . RESULTS OF EXAMINATIONS AND TESTS

3 . 1 Visual and Macroscopic Examination

Every blade of the compressor turbine disk and blade

assembly had been fractured, with the possible exception of

two missing blades which had been attached to the fractured

fir tree part of the disk. A number of the fractured blades

- 1 -

can be seen in figures 1 and 3. In general, the fractures

occurred so that the outermost part of the blades remaining

in the assembly ranged from about 1/16 to 1/4 inch beyond the

top of the blade base (arrows B, figures 1 and 3) . One blade

(arrow A, figure 1) had fractured near the base of the disk

fir trees. This was in the region of the disk where three of

the disk fir trees had fractured. A few of the blade fracture

surfaces exhibited rather severe mechanical damage (arrow C,

figures 1 and 3) , apparently having been caused by relative

circumferential motion between the blades and either another

compressor part or a foreign object. Except for those which

had suffered mechanical damage, the fracture surfaces of the

blades were dark gray in color.

Both sides of the disk were discolored. The portion

between the ring containing the balancing weights and the disk

hub ranged from light blue near the balancing weight ring to

no apparent discoloration at the hub. Discoloration in this

area was greater on the side of the disk opposite to that

containing the spline. The regions between the balancing

weight ring and the outside of the disk were discolored about

equally on both sides. The coloring ranged from light blue

near the ring through straw, violet, green, and back to straw.

The fir tree regions of the disk and the blades were dark

gray in color.

About 1/3 of the circumference of the outside of the

balancing weight ring on the side of the disk containing the

spline had been roughened by circumferential rubbing. Part

of this damage can be seen in figure 1 (arrow D)

.

The heads of all the rivets and the sides of all the disk

fir trees on the side of the disk containing the spline had

suffered mechanical damage by apparent circumferential rubbing.

Many of the blade fir trees also suffered similar damage.

Examples of this fir tree and rivet damage are shown in figure

2. The severest damage to the sides of the disk fir trees was

found in about 1/3 of the disk circumference. About 2/3 of

the more severely damaged area was in the same 1/3 of the disk

which contained the damage to the balancing weight ring

discussed above. Many of the blade bases were askew as can be

seen in figure 3. This was probably caused either by the

rubbing action or contact of the blades with another component

or foreign object. Some of the blade bases were somewhat loose

on the fir tree assemblies and could be moved slightly.

- 2 -

There was a line of fretting damage which traversed

between 1/4 and 1/3 of the circumference of the disk hub

between the spline and the end of the hub (figure 4) . This

damage was located essentially opposite to the damage to the

balancing weight ring.

Every disk fir tree was cracked in at least two places,

and possibly three. An example of these cracked regions is

shown in figure 5. One of the cracked regions, indicated by

A, is in the first root from the top of the fir tree. The

second cracked region, indicated by B, figure 5, is on the

opposite side of the fir tree in the second root from the top.

Many, and perhaps all, of the fir trees were cracked in a

third region designated C in figure 5. There was considerable

mechanical damage in this area of the trees and it was not

possible to determine in all cases whether a crack was

present without some preparation of the surface (which did

not appear to be warranted) . Cracks in other areas of some

of the disk fir trees were noted. In the tree shown in

figure 5, the cracks in region B appear to be propagating

toward those in regions A and C, and conversely, cracks in

regions A and C appear to be propagating toward those in

region B.

Three of the disk fir trees had fractured completely.

The fracture surfaces of these fir trees are shown in figure

6. Fracture 1 had occurred across the tree through region

C-C, fracture 2 through region BC, and fracture 3 through

region A-A' , figure 5.

3 . 2 Fractographic Examination

The fracture surfaces of the three fractured fir tree

sections of the disk were blue in color in some areas

(primarily near the centers) and dark gray in others

(primarily near the leading and trailing edges as the disk

would rotate in service) . The three fracture surfaces were

examined with the scanning electron microscope (SEM) . Before

the examination, the surfaces were cleaned in an ultrasonic

bath with water and detergent. This cleaning treatment did

not remove the discoloration from the surfaces indicating

that it v/as apparently not of a superficial nature.

The SEF examination revealed intergranular cracking as

the principal mode of fracture in the apparent crack initiation

areas on both leading and trailing edges (as the disk would

rotate in service) of fractures 1 (location C-C t figure 5)

and 2 (location BC, figure 5) and on one side of fracture 3

(location A-A', figure 5). The cracks propagated in an

- 3 -

intergranular mode for varying distances up to one half the

width of the fir tree before the fracture was completed by

ductile overload, as evidenced by dimpled rupture as the

principal fracture feature. Representative examples of the

intergranular cracking are shown in figures 7 and 3, and a representative example of dimpled rupture in an overload

region is shown in figure 9. Figure 8 appears to exhibit

some corrosion product. The dark gray areas of the fracture

surfaces exhibited the intergranular cracking and the blue

areas exhibited dimpled rupture.

3 . 3 Hardness ^Measurements

Knoop microhardness measurem.ents (KHN) at a load of 500

grams were made on a circumferential section through some of

the cracked disk fir trees. The KHN500 hardness ranged from

352 to 394. The approximate Brinell equivalent hardness

values ranged from BHN 332 to 362 which is well within the

requirements of United Aircraft of Canada in-house specification CPW 203 which called for a Brinell hardness within the

range of 313 to 403, or equivalent.

3 . 4 Metallographic Examination

A circumferential section through several of the disk

fir trees was examined metallographically . A low magnification photograph showing parts of the fir tree portions of

two blades (top left and top right in the figure) , parts of

two rivet heads, and parts of two of the disk fir trees is

shown in figure 10 v/ith the specimen unetched. Several

cracks can be seen, including the cracks in regions A, B,

and C, figure 5. These crack regions are again indicated by

A, B, and C in figure 10. The dark areas in the photomacrograph are filled with mounting material, hence the only

parts of the section that show are those in the plane that

has been polished. The area containing crack regions A and

B is shown at higher magnification in figure 11. There is

considerable deformation in the region between the crack

regions, and it can be seen that the cracks in the two regions

appear to be propagating toward each other.

A higher magnification photomicrograph of one of the

cracked regions (figure 12) shows that the cracks are intergranular in nature as indicated by the f ractographic examination of the fractured fir trees. The microstructure

consists of carbide particles in a gamma matrix and appears

to be free of stringers. A fairly thin region near the surface

of the tree is outlined in figure 12. This is a zone

- 4 -

apparently caused by reaction of the material of the disk

with the service atmosphere at elevated temperatures, and is

referred to by the manufacturer as a depleted zone. This

depleted zone was found along the walls of some of the

cracks in some places, an example of which is shown in figure

13.

4 . DISCUSSION OF RESULTS

The discoloration of parts of the compressor turbine

disk assembly was apparently caused by subjecting the assembly

to high temperatures, but since the assembly is expected to

operate at elevated temperatures, the discoloration is likely

a consequence of normal operation.

The fractured blades and disk fir trees, the fretted hub,

and the rubbing damage to the sides of the disk fir trees,

the sides of the blade fir trees, the rivet heads, and the

balancing weight ring indicate that the disk assembly was in

contact v/ith either another component (s) of the engine or a foreign object. It may be possible that some of the fractured

blade fragments caused some of the damage to the sides of the

disk fir trees. The cracks in the fir tree areas of the disk

may have been caused by stresses transmitted to them through

the blades when the blades were in contact with other components or a foreign object, or by overstressing in operation.

Both f ractographic and metallographic examinations

indicated that the cracks in the fir trees propagated

initially in an intergranular mode. For the three disk fir

trees which had completely fractured the final rupture

occurred in ductile overload. The discoloration of the

fracture surfaces indicates that the assembly was probably

subjected to high temperatures after the cracks had

propagated to fracture, although the speed at which the

cracks propagated is not known.

The depleted zone which appeared along the surface of

the fir tree portions of the disk (and perhaps elsewhere at

the disk surface) was apparently caused by a reaction of the

disk material to the service atmosphere. The fact that the

depleted zone appeared along some of the fir tree cracks

indicated that these cracks had been in existence with the

assembly in service long enough for the zone to form.

Nothing unusual was noted in the microstructure of the

disk material. The microstructure consisted of carbide

particles in a gamma matrix, which is normal for Waspaloy

in the heat treated condition called for in United Aircraft

of Canada Specification CPW 203.

- 5 -

The results of the hardness measurements indicated that

the disk material met the hardness requirements as set forth

in United Aircraft of Canada Specification CPW 203..

5 . CONCLUSIONS

1. The extensive damage to the compressor turbine disk and

blade assembly was apparently caused by contact of the

assembly with other components of the system or with a foreign object.

2. The discoloration of the disk surface appeared to be

due to exposure to high temperatures and is likely a consequence of normal operation.

3. Cracks were found in all the disk fir trees. They

probably were caused by stresses transmitted to them

through the blades when the blades were in contact

with other system components or a foreign object.

4. The fractures in the three completely fractured disk

fir trees initiated and propagated some distance in an

intergranular mode before final failure occurred in

ductile overload.

5. A depleted zone along some of the disk fir tree crack

surfaces indicates that the assembly was probably in

the service atmosphere for some time after those

cracks formed. A depleted zone was also found at the

surfaces of the disk fir trees.

6. The microstructure appeared normal for Waspaloy in the

prescribed heat treatment.

7. The hardness of the disk material met United Aircraft

of Canada Specification CPW 203.

6. ACKNOWLEDGEMENT

Mr. L. C. Smith of the NBS Mechanical Properties Section

prepared the specimens for the metallographic and fractographic examinations and performed the hardness measurements.

- 6 -

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Figure 2. Part of the side of the turbine disk and blade assembly

containing the spline showing circumferential rubbing

damage to the rivet heads and to the fir tree areas of

the disk and blades. X 4

Figure 3. Part of the compressor turbine disk and blade assembly

showing blade base askewness (arrows E) and the fracture

surfaces of several blades. Arrows B indicate fractures

representative of most of the blades. Arrow C indicates

a blade fracture surface exhibiting severe mechanical

damage. X 4

Figure 4. Part of the compressor turbine disk hub between

the spline and the end of the hub showing fretting

damage (indicated by arrow) . X 10

One of the fir trees of the disk (center) and

two blade fir trees. The regions of crack

initiation in the disk fir trees are indicated

by A, B, and C. Of the three fir trees which

had fractured, the fracture crack passed

through A-A', in one, through B-C in one, and

through C-C for the other. X 10

I Figure 5.

I

I

Figure 6. Fracture surfaces of the three fractured disk

fir trees, indicated by arrows 1, 2, and 3.

Part of one of the blade fir trees which had

fractured and remained attached to the disk

is indicated by the unmarked arrow. The

locations of the scanning electron photomicrographs shown in figures 7, 8, and 9 are

indicated by A, B, and C, respectively. X 4

Figure 7 . Scanning electron photomicrograph showing

intergranular cracking. The area shown

was located approximately at A on fracture

1, figure 6. This is representative of the

fracture features exhibited by the dark gray

areas of the fracture surfaces. X 520

Figure 8. Scanning electron photomicrograph showing

intergranular cracking. There appears to

be some corrosion product here. The area

shown was located approximately at B on

fracture 3, figure 6. X 330

Figure 9. Scanning electron photomicrograph showing

dimpled rupture . The area shown was located

approximately at C on fracture 3, figure 6.

This is representative of the fracture

features exhibited by the blue areas of the

fracture surfaces. X 850

Figure 10. Circumferential section through assembly

showing parts of two blade fir trees (upper

right and upper left) , parts of two rivet

heads, and parts of two disk fir trees.

Crack regions A, B, and C are indicated as

in figure 5. Dark areas contain mounting

material. Unetched. X 7

Figure 11. Crack regions A and B from figure 10 at

higher magnification. Section was etched

to show the deformation in the material

between the two cracked regions

.

Etch: Kalling's. X 40

Figure 12. Etched circumferential section in crack

region B of one of the disk fir trees.

The intergranular nature of the cracks is

evident. A narrow region near the surface

of the tree (indicated by arrows) termed a depleted zone can be seen.

Etch: Kalling's. X 200

Figure 13. Etched circumferential section showing

depleted zone along part of the crack

wall as well as along the surface of

the tree

.

Etch: Kalling's. X 500

us. C •_ "

- ;

L

'.' M. BI3LICG^ A. -IC DATA

1. !M m.li :\ I lOX OR KIJ'OKl' \(.). NBSIR 73-2842. u.iv't Acot in 3. Kcci in cut' <; Acc<;s>uin .\ i

4. I \' I' exm'ilnation of united aircraft of canada

co::pressor turbine disk and blade assej^ibly.

5. I 'ub 1 u > t ion D.itc October, 1973

6. Hi . 1 ::. la ^ Or ijani.- .![ u.ii

7. Al'THOR.S)

T. Robert Shives

8. l'crIi>r:;>inL; Org,in;:Mt;vm

NBSIR 73-284

9. PERFORMiXe, ORG \\ 1/. A ITON XAMl- AND AODRESS

NATIONAL HLRFAU 01- STANDARDS

DEPARTMKNT 01- C0MMI-:RCE

WASHINGTON, D.C. 20234

10. Project, I'ask, Work l.'nit No. 3120412

11. Contract/ Grant No.

12. Sponsoring Organization Name and Address Bureau of Aviation Safety

National Transportation Safety Board

Department of Transportation

Washington, D. C. 20591

13. Type of Report & Period

Covered _ . t _ Failure

Analysis Report

14, Sponsoring Agency Code

15. SUPPLEMENTARY NOTES

16. ABSTRACT (A 200-word or less factual summary of most significant information. If document includes a significa-c

bibliography or literature survey, mention it here.) At the request of the Bureau of Aviation Safety, Department of

Transportation, a failed compressor turbine disk and blade assembly from

a United Aircraft of Canada aircraft engine was examined by the NBS

Mechanical Properties Section. Every blade still connected to the

assembly when it was subm.itted for examination was fractured and every

disk fir tree which had not completely fractured was cracked in at

least two and possibly more places. Three of the disk fir trees had

fractured completely . A fractographic examination revealed that the

cracks which lead to fracture had originated and propagated for some distance in an intergranular mode before final failure occurred in

ductile overload. The microstructure of the material appeared to be normal for the specified treatment and the hardness met specifications.

17. KEY UORDS (Alphabetical order, separated by semicolons)

Compressor turbine disk; Intergranular cracking; Turbine blades;

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