Synopsis
- 10 -
Examination of the fracture using a scanning electron microscope rycle fatiaue p r o p a m n (SEM) disclosed fracture features typical of h' from these origin areas. Fatigue had propagat2down and through the bottom of the fitting and out. (See Figure 4 , Appendix F.) Approximately 40 percent of the fractured cross sectional area of the fitting appeared cracked. The remaining sections of fracture away from this fatigue region were typical of overload separations.
~ ~ ~ ~~~~ - ~~
Some areas were missing along the fracture line that corresponded to the overload portion of the separation. These missing areas suggested that the fitting opposite the fatigue region and near its forward side had broken in compression since compression breaks tend to fragment the fracture.
The missing area on the forward side was seen by comparing the mating fracture halves and noting that the fracture on the inboard half went through the small hole used to locate the tube in the fitting while on the outboard half the fracture in the same area progressed through the bolt hole. (See Figure 3 , Appendix F.)
Detailed examination of the fracture origin area disclosed two origin sites, both of which appeared to be at discontinuities in the radius (See Figures 5 and 6, Appendix F ) . The two origins produced two planes of fatigue fracture slightly offset from one another. The two fatigue cracks grew into one large crack a short distance out from the radius. Because of the extent and symmetry of the crack arrest markings and the locations of the origins with respect to a ratchet mark (step- one origin like portion of the fracture connecting the fatigue planes), was considered secondary and the other was considered the main origin.
A substance was found which completely filled the discontinuity at the secondary origin and partially filled the discontinuity at the main origin. (See Figures 5 and 6, Appendix F.) Energy dispersive X - ray analysis of the substance at the main origin area disclosed elements normally associated with the fitting alloy system along with an appreciable amount of sulfur and small amounts of calcium and potassium all of which are foreign elements. The silicon energy peak was strong indicating silicon as a primary element of the substance; the fitting alloy normally contains less than 0 . 4 by percent weight of silicon.
The radius in the origin area was mostly covered by an adhesive (EC - 2214) used during fabrication of the assembly. Analysis of this adhesive material indicated that it was high in aluminum and silicon with some sulfur, chlorine, potassium and calcium.
Hardness and electrical conductivity measurements of the fitting gave values averaging Rockwell "B" 83 and 40 percent International Annealed Copper Standard, respectively. These measurements, as well as the microstructural characteristics, appeared normal for 7075 aluminum alloy heat treated to the T73 condition (specified material and heat treatment).
Investigation
History of the Flight
- 11 -
Examination of the right - hand aft main landing gear fitting and tube assembly showed fractures that were all typical of an overload separation. No evidence of fatigue or other type of pre - existent cracking was found on this assembly.
1.17 Additional Information 1.17.1 History of Heliport Operation Heliport operations began in 1965 from the Pan Am Building with New York Airways operating Boeing Vertol 107 equipment. There was public pressure against the NYA petition because of noise and safety reasons. Hearings were held and the operation was approved. A 5 - year permit from the NYC Planning Commission was issued for the period 1964 through 1969. The NYC Department of Marine and Aviation then issued an Air Facility Permit for 1 year, which was later renewed for another 1- year term.
In February 1968, NYA ceased operations from the heliport because of a contractual disagreement between Pan American World Airways and Trans World Airlines. During the disagreement, the facility permit for the heliport expired. Shortly before the expiration date NYA requested a renewal. The Department of Marine and Aviation refused to renew the permit because (1) NYA was not operating onto the roof heliport currently, (2) NYA presented no plans in the renewal petition to begin operations at a set date, ( 3 ) public pressure was such that a hearing should be held. Since NYA was not operating to the roof, the city determined that there was no point in renewing the permit at that time. No further petitions were presented, and in 1969 the NYC Planning Commission permit expired.
On November 24, 1976, the New York City Planning Commission held a public hearing to consider the application of New York Airways, Inc., for a resumption of scheduled helicopter operations from the roof top heliport on the Pan American Building. Before the hearing, local community planning boards 5 and 6, representing the neighborhoods most directly affected by the proposed operation, gave their approval of the granting of an operational permit after an extensive review of the facility. The operation was then approved by the New York City Board of Estimate, the Department of Marine and Aviation, and by the Federal Aviation Administration.
It The resolution by the Board of Estimate, City of New York, to grant a permit to operate the Pan American Building Heliport was approved December 6, 1976, and adopted January 20, 1977. The special permit was for a 3 - year period.
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On January 27, 1977, New York Airways, Inc., applied to the Department of Marine and Aviation for an Air Facility License for the operation of a heliport on the roof of the Pan American Building. This license was granted on January 31, 1977, for 1 year. Operations began on February 1, 1977, and were conducted without incident until this accident. During this period New York Airways conducted 7,240 helicopter operations from the rooftop heliport.
1.17.2 History of S - 61L Forward Lower Landing Gear to Fuselage Attachment Fitting Failures On July 15, 1963, a forward lower landing gear to fuselage attachment fitting failed when a Los Angeles Airways S-61L, N300Y, was parked with rotors turning at the American Airlines gate area, Los Angeles International Airport. While ground personnel loaded mail aboard the aircraft, a snapping noise was heard, and the helicopter tipped to the right and rolled over on its right side. Outboard sections of the five main rotor blades contacted the ramp surface and separated.
One person on the ground was injured and windows in the airline terminal building were broken.
Investigation revealed that forward lower landing gear fitting, P/N S6125-50312-22, had failed in fatigue and had separated, allowing the right main landing gear to collapse. This part had a total time of 1912:43 hours.
As a result of this occurrence, the manufacturer redesigned the forward and aft fittings, both left and right, as follows: 1. Forging material was changed from 7079T6 aluminum alloy to 7075T73 aluminum alloy.
2 . Wall thickness of the tube portion of the fitting was increased from .120 to .160 inches.
3. The internally machined radius was increased to .12 inches.
The redesigned fittings were installed on all existing S - 61L models (four at that time) per Sikorsky Service Bulletin 61 B25 - 1 dated September 18, 1963. Appropriate engineering drawings were changed requiring installation of the improved fittings. (PN's 6125 - 50333 - 21 - 22 and PN's 4125-50334-21-22).
Subsequently, all S - 61L model aircraft, including those purchased by New York Airways, Inc., had the improved fittings installed. This accident resulted from the first failure since the redesigned fittings had been installed.
- 13 -
1.17.3 Normal Helicopter Operating Procedures Experience has shown t h a t t h e o v e r a l l s a f e t y of h e l i c o p t e r operations is enhanced by o p e r a t i n g t h e engines and r o t o r system during t h e f r e q u e n t passenger enplaning and deplaning of normal scheduled operations. Dynamic components of a h e l i c o p t e r ' s main r o t o r system a r e n o t designed f o r f r e q u e n t s t o p s ; t h e r e f o r e , r o t o r shutdown at each s t o p would s h o r t e n t h e s e r v i c e l i f e of t h e s e components which are critical t o s a f e t y of f l i g h t and would i n c r e a s e t h e p o s s i b i l i t y of component f a t i g u e .
The s t a b i l i t y of a h e l i c o p t e r is increased when t h e r o t o r b l a d e s are t u r n i n g during high and v a r i a b l e wind conditions. Also, b l a d e f l a p p i n g a t low r o t o r RPM during such wind c o n d i t i o n s i n c r e a s e d t h e p o s s i b i l i t y of b l a d e damage t o t h e a i r c r a f t ' s airframe.
1.18 New I n v e s t i g a t i o n Techniques None 2. ANALYSIS The a i r c r a f t was c e r t i f i c a t e d , equipped, and maintained according t o a p p l i c a b l e r e g u l a t i o n s . The a i r c r a f t ' s powerplants and systems were not f a c t o r s i n t h e accident. The g r o s s weight and c.g. were w i t h i n p r e s c r i b e d limits. The f l i g h t c r e w was properly c e r t i f i c a t e d and each crewmember had received t h e t r a i n i n g and off - duty time p r e s c r i b e d by a p p l i c a b l e r e g u l a t i o n s .
t o t h e f l i g h t c r e w , According passengers, and o t h e r w i t n e s s e s , t h e landing on t h e h e l i p o r t The a i r c r a f t w a s t a x i e d t o t h e was g e n t l e .
normal p o s i t i o n without i n c i d e n t , and approved procedures were followed f o r passenger operations. The f a c t t h a t t h e engines were o p e r a t i n g , and t h e r o t o r was t u r n i n g had no b e a r i n g on t h e f a i l u r e of t h e landing gear.
N o c o n t r o l input was made which could have overloaded t h e p a r t which f a i l e d .
The c a p t a i n s t a t e d t h a t h e had h i s hands on t h e c o n t r o l s , and t h a t t h e c o l l e c t i v e was bottomed, which would produce minimum torque on t h e engines. This s e t t i n g was v e r i f i e d by t h e f i r s t o f f i c e r . The c y c l i c was i n t h e n e u t r a l p o s i t i o n according t o t h e c a p t a i n and t h e f i r s t o f f i c e r . The f i r s t o f f i c e r s t a t e d t h a t as t h e a i r c r a f t began t o r o l l , he saw t h e b l a d e t i p p a t h r o t a t e w i t h t h e a i r c r a f t and remain c o n s t a n t i n t h e windshield i n r e l a t i o n , t o t h e f u s e l a g e . The c o n s t a n t r e l a t l o n s h i p shows t h a t no c y c l i c i n p u t was induced which i n t u r n would have t i l t e d t h e rotorhead and t h e b l a d e t i p p a t h plane.
The S a f e t y Board reviewed t h e a c t i o n s taken by t h e crew from t h e f i r s t sound of f a i l u r e u n t i l t h e b l a d e s s t r u c k t h e h e l i p o r t surface,, and concluded t h a t t h e a c t i o n s taken by t h e crew -- to c l o s e t h e ESC's--' were c o r r e c t . The l o c a t i o n of t h e a i r c r a f t and t h e number of people \,
- 14 -
around the helicopter, coupled with an unexpected noise from an area \ initially believed by the captain to be the rotor system, dictated that \ the engines be shut down immediately. The fact that both ESC's were !
shut down so quickly probably prevented further damage and injury and possibly prevented a fire.
The pilot's reaction of not applying the rotor brake was also correct. The rotor brake is not designed to stop a rotor head turning at 100 percent NR quickly. It is designed to stop the rotor blades once NR is below 40 percent. For an emergency shutdown the lever may be forced forward into the full on position, after closing the engine W M control with a delay time of 5 seconds. Since the captain barely had time to close the ESC's before the blades struck the ground, the Safety Board concludes that he could not have used the rotor brake for an emergency stop of the rotor blades. At 100 percent NR, the application of the rotor brake will cause the brake to heat up and possibly burn out, thereby creating a fire hazard.
The Safety Board, therefore, concludes that the failure of the gear was not the result of a pilot input.
The fracture of the right hand landing gear forward fitting, P/N S6125-5033-22, stemmed from a fatigue crack through 40 percent of the cross section.
The crack location in the bottom portion of the fitting suggested that bending stresses, which compressed the top of the fitting tension on the bottom, had caused the fracture to begin and and placed propagate. Planar orientation of the crack was diagonally downward and inboard through the support rib between the fuselage and lower diagonal tube attachment. Since fatigue cracks tend to propagate in a plane perpendicular to the direction of principle tension stress, a tension force along the diagonal tube may also have contributed to the crack propagation.
The material properties of the fitting, including the chemical composition, were considered normal for 7075-T73. Originally the fitting was designed using the alloy 7079-T6 (P/N S6125-50312). However, problems with stress corrosion cracking warranted changing the material to 7075- T73 (T73 condition is highly resistant to stress corrosion).
The radius showed no appreciable attack from corrosion, indicating this area vas adequately protected against it. There was no indication of stress corrosion on the fracture surface although pitting of the fracture was noted. This pitting, however, was much more intense in the latter stages of fatigue propagation, indicating that the corrosion medium may have entered when the crack opened to the atmosphere (breaking through the support rib and lower portion of the fitting).
Injuries to Persons
Damage to Aircraft
Other Damage
Personnel Information
Aircraft Information
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Two holes, however, did enter into the interior of the tube.
These holes were located along the inboard edge of the adhesive just inboard of the tube end and were open between the external and internal radius of the fitting. An access, therefore, existed for water or other mediums to penetrate into the interior and collect along the lower portion of the radius. If the adhesive in the radius area contained minute holes through its thickness, then corrosion attack would have been highly localized, and would have produced discontinuities such as those found on the failed fitting.
The discontinuities found could also have been produced by normal microconstituent particles or phases in the material which were exposed to the surface. The sizes of the discontinuities, however, were much larger than the compound phases or particles found in the microsections which would mean that the discontinuities would have to be unusually large particles or groupings of second phase particles.
The substance found in the cavities of the discontinuities is believed to be the adhesive used in the assembly. It is not known whether the adhesive entered the cavity before or after the fracture.
The Safety Board, therefore, concludes that the fracture of the right main landing gear was the result of fatigue originating from a small surface pit of undetermined source.
Although, in this case, the rotating main rotor blades resulted in fatalities when the right landing gear failed, there is a good possi - bility that similar fatalities could have occurred if gear failure had occurred during shutdown or starting procedures. In considering all of the safety aspects of continued rotor rotation during enplaning and deplaning operations versus frequent rotor shutdowns, the Safety Board concludes that continued rotor rotation during such operations is, in most cases, safer. This conclusion is based on the fact that, (1) frequent rotor shutdowns shorten the service life of the rotor components and the engine by introducing more fatigue cycles thereby reducing the overall safety of operations, (2) potential safety hazards such as rotor brake fires, engine acceleration malfunctions, and main blade to fuselage contact due to excessive flapping at low rotor R P M are introduced by repeated rotor brake applications, and ( 3 ) the stability of the helicopter is increased due to the gyroscopic effects resulting from the rotating blades. Continued rotation is especially important for safety during variable, high wind conditions.
* The heliport was certificated properly under 14 CFR 139. All the equipment which was required to meet the crash fire rescue (CFR) criteria was met or exceeded. The personnel requirements, as stated in the heliport manual, were met or exceeded.
Meteorological Information
Communications
Aerodrome Information
Flight Recorders
- 16 -
The Safety Board also concludes that, although some confusion existed and some misunderstanding of CFR responsibilities were apparent, the CFR activities were effective for the situation.
The Safety Board recommends that the cockpit cabin door of the aircraft remain open. Helicopters operated under 14 CFR 127 are not required to have the cockpit door closed and locked as are aircraft operated under 14 CFR 121. Since the door is not usually used in normal operations, the Safety Board believes that it is unnecessary and could become an obstruction between the cockpit and cabin during emergency conditions.
3. CONCLUSIONS 3 . 1 Findings 1. The aircraft was certificated and maintained according to approved procedures.
2. All crewmembers were certificated and qualified for the flight.
3. The airport was properly certificated under 14 CFR 139, without exemptions.
4 . The fracture of the right main landing gear forward fitting resulted from a fatigue crack.
5. The fatigue crack had initiated along the 0.12 inch internal radius near the bottom of one of the two loca - tion holes where the hole changes from a cylindrical to a conical shape.
6. Fracture features were typical of high cycle fatigue propagation from the origin areas and propagation down and through the bottom of the fitting.
7. All fractures outside the vicinity of the fatigue origin regions resulted from overload separations.
8. Hardness and electrical conductivity measurements of the failed fitting gave values normal for 7075 aluminum alloy 1 heat treated to the T 7 3 condition.
9 . The failure of the landing gear did not result from any pilot inputs or operational overloads.
10. The flightcrew's decision to close the EX'S after a malfunction was suspected was proper.
Wreckage and Impact Information
General Examination
Landing Gear Examination
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The rotor brake could not have been used effectively to 1 1 .
halt the rotor system in the short time interval between failure of the landing gear and rotor blade impact.
12. The CFR procedures for the heliport were adequate for the emergency.
13. New York Airways personnel accomplished effectively the CFR duties, although there was some confusion concerning overall supervision of specific duties.
3.2 Probable Cause The National Transportation Safety Board determines that the probable cause of the accident was the fatigue failure of the upper right forward fitting of the right main landing gear tube assembly.
Fatigue originated from a small surface pit of undetermined source. All fatalities were caused by the operating rotor blades as the result of the collapse of the landing gear.
4 . SAFETY RECOMMENDATIONS As a result of this accident, on May 18, 1977, the National Transportation Safety Board issued the following recommendations to the Federal Aviation Administration: “Issue an Airworthiness Directive to require an immediate one - time inspection by an approved method on both the forward and aft main landing gear attachment fittings, right and left, on all Sikorsky Model 61L series helicopters having similar
installations. (Class I - Urgent Followup) (A - 77 - 32)
a , Reevaluate the current inspection interval and issue requirements for more frequent periodic inspections if necessary to insure continued safe operation. The inspection interval could be based on a set number of operating cycles instead of
an established operating time. (Class I1 - Priority Followup)
(A - 77 - 33)” Upon receipt of these recommendations, the Federal Aviation Administration issued a telegraphic airworthiness directive which: ’ . (1) Required, prior to next flight, a fluorescent penetrant j inspection of the forward and aft main landing gear attachment fittings, right and left, on all affected ~.
Sikorsky Model 61 series helicopters. In addition, a visual inspection was required prior to the first flight of each day.
Medical and Pathological Information
Fire
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(2) Required the reevaluation of the current inspection interval and issue requirements for more frequent periodic inspections if necessary to insure continued safe operation.
The inspection interval could be based on a set number of operating cycles instead of an established operating time .
As a result of the cockpit door of the S - 61L sliding almost closed and jamming, on July 13, 1977, the Safety Board subsequently recommended that the Federal Aviation Administration: " Require that the sliding cockpit door on the Sikorsky S - 61L helicopter be removed or retained open so that it cannot obstruct the entrance from the cockpit to the cabin area.
(Class 11 - Priority Followup) (A - 77 - 51) BY THE NATIONAL TRANSPORTATION SAFETY BOARD / s / KAY BAILEY Acting Chairman / s / FRANCIS H. McADAMS Member / s / PHILIP A. HOGUE Member / s / WILLIAM R. HALEY Member October 13, 1977
Survival Aspects
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APPENDIX A INVESTIGATION AND HEARING 1 . Investigation The Safety Board was notified of the accident about 1745 on May 16, 1977. The investigation team went immediately to the scene.
Working groups were established for operations/weather/airports/air traffic control, human factors, witnesses, structures/systems, powerplantsl maintenance records, and cockpit voice recorder.
Participants in the on - scene investigation included representa - tives of the Federal Aviation Administration, Sikorsky Aircraft, New York Airways, Inc., New York City Transportation Department, Air Lines Pilots Association, and Association of Flight Attendants.
2 . Hearing A public hearing was not held.
Tests and Research
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APPENDIX B PERSONNEL INFORMATION Captain Lee G. Richmond Captain Lee G . Richmond, 46, was employed by NYA on February 24, 1964, as a first officer. He was qualified initially in the S61N on March 22, 1964, and upgraded to Captain on November 17, 1970. His initial Category - A Edge qualification was on April 23, 1971.
Captain Richmond's last line check was on June 28, 1976. He had completed proficiency checks/recurrent training on January 12, 1977, and July 1 3 , 1976. He received an Edge Procedure Checkout and Line Check on February 1, 1977. This period also included a 1 hour ground school on edge procedures.
Captain Richmond holds Airlines Transport Pilot Certificate
No. 644892 dated November 11, 1970, with the following ratings: Rotorcraft -
Helicopter BV - 107 - 11 (VFR only) SK - 61; Commercial Privileges airplane single - engine, land and sea; airplane multi - engine land, glider, instruments.
His first - class medical certificate was dated April 21, 1977.
It had no limitations.
Captain Richmond had about 11,721 total hours of flight'time at the time of the accident, 9,000 of which were in helicopters. He had about 2,200 hours in S61 helicopters. In the previous 30 days he had recorded 50:30 hours as a S61 helicopter captain and 9:20 hours other helicopter time.
In the 24 - hours before the accident he had recorded 3:48 hours rotor time as an S61 captain.
First Officer John F. Flanagan First Officer (F/O) John F. Flanagan, 31, was hired by New York Airways on April 8, 1977. His initial 561 qualifications was on April 16, 1977, and his Edge Procedure Qualification was completed on April 15, 1977.
First Officer Flanagan holds Commercial Certificate No. 1987361 with the foJlowing ratings: Airplane single - engine land, rotorcraft- helicopter instruments including helicopter. The date of his certificate was January 6, 1970. His second - class medical certificate was dated August 26, 1976, and had no limitations.
Additional Information
History of Heliport Operation
- 21 - APPENDIX B
First Officer Flanagan had a total of 1,768.4 hours of flying time, with 1,339.2 hours recorded in helicopters. His total 561 time was 61 hours, all flown in the 30 days prior to the accident. He had recorded 3:48 hours rotor time in the 24 hours before the accident.
The day of the accident was the fourth consecutive day that each pilot had flown. On the first day of this sequence they did not fly together. However, they flew together on the second and third days, and the day of the accident. They reported for duty at JFK at 1402 on the day of the accident. Each had been off duty since 2117 the previous day. At the time of the accident each pilot had been on duty 3:33 hours.
Flight Attendant Lammie Chevalier Flight Attendant Chevalier was employed by New York Airways in 1973. She was current and qualified to perform her prescribed duties.
Fuselage Attachment Fitting Failures
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APPENDIX C AIRCRAFT INFORMATION Sikorsky S-61L, N619PA (SIN 61427) was manufactured in June 1968. It was owned by the General Electric Credit Corporation of Georgia and operated by New York Airways, Inc. It was certificated and maintained according to procedures approved by the FAA. At the time of the accident the aircraft had accumulated about 6,913:15 flight hours and 7:22 flight hours since its last major inspection.
The aircraft was equipped with two General Electric CT58-140-2 engines. Engine serial numbers and times are as follows: Position Serial No. Total Time 1 295063C 7,201:44 2 295069C 6,517:52 I UNION CARBIDE CHEMICAL BANK BLDG BUILDING I I I U I
ST @
E 47
AMERICAN BRAND!
BLDG
E 46
NEWYORK 466 GENERAL LEXINGTON
>
BUILDING BLDG
E
P,
Q
z L
I
L E 45
p y - ACCIDENT
POST OFFICE
-
E 4 4
GRAYBAI HOTEL BUILDIN( BILTMORE GRAND CENTRAL TERMINAL ‘331
I MADISON
I AV
ST
E 42
I Site Co-ordinates: Long. 7 3 O 58’ West
Lat. 40° 45‘ North
APPENDIX ' 0 LEGEND 1. C-5" Outboard Section - Main Blade Spar 2. 4 ' 4 " Spar Section - Outboard End Tip 3. C-1'' Spar Section - Outboad End Tip 4. 4-0" Spar Section - Outboard End Tip 5. 0-11'' Spar Section - Outboard End Tip 6. 2-3" Spar Section - Outboard End Tip 7. C-2" Outboard Section - Main Blade Spar 8 . 2 Twelve Inch Wide Trailing Edge PoCketO 9. 10-6" Center Section - Main Blade Spar IPenetrated Window of Office 36th Floor) 10. 2-4" Outboard Section - Main Blade Spar 11. 1 Twelve Inch Wide Trailing Edge Pocket 12. 3'W Outboard Section - Main Blade Spar Ion Heliportl
I
13. 4 " Z Outboard Section - Main Blade Spar FERICAN BRANDS [Outside Room 369-466 Lexington Ave.1 14. 1 9 - 2 Center Section - Main Blade Spar BLDG Ion Heliport Just Forward of Aircraft Nose) 15. Miscellaneous Blade Parts - PocketO.
I Leading Edge Counterweights 111 Locations)
c
e Helicopter Location
POST @
A
OFFICE 8
I I I CENTRAL GRAYBAR TERMINAL BUILDING I I NATIONAL TRANSPORTATION SAFETY BOARD WASHINGTON, D.C.
Wreckage Distribution Chart New York Airways Inc.
Sikorsky S61L N619PA Pan Am Building Rooftop Heliport I Manhattan, New York May 16.1977
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APPENDIX E FIGURE 1 APPENDIX E FIGURE 2
- 26 -
ENERGY ABSORBING STRUT 6125 - 50301 AFT Flll'lNG 6125-50334-12 I
I I / / / / FITTIN G ASSEMBLY
6125-50317-3
L CLY
\
OLEO STRUT FWD 6125-50302 \ S - 61L MAIN LANDING GEAR - 27 - APPENDIX F Figure 1. Overall view of the components, as received.
1.) Forward main landing gear fitting and tube assembly.
2 . ) Aft main landing gear inboard fitting with portion of tube.
Figure 2. Right hand main landing gear fitting (P/N 56125 - 50333 - 2) with the fractures (arrows " a " figure 1) placed relative to each other as if intact. Inboard attachment bolt was removed from the hole.
APPENDIX F - 28 -
Looking Outboard Looking Inboard Figure 3 . Mating fracture surfaces on the P/N S6125-50333-2 fitting. Both photographs approximately X1 1 / 4 Figure 4 . Portion of inboard fracture surface removed for detailed fractographic examination. Arrow " 0 " denotes main origin and arrow "S" locates secondary origin while dashed lines indicate approximate extent of fatigue propagation from these origins.
- 29 - APPENDIX F Figure 5. SEM photograph showing secondary origin bracket
,, , e
S (see arrow "S" in figure 4 for location). Fracture surface is above origin and hole radius is below. Arrow- heads depict fatigue propagation direction. X100 .
Figure 6. SEM photograph of main fatigue origin area (see arrow " 0 " figure 4 for location). Arrowheads indicate fatigue propagation directions. X90 - - 30 - APPENDIX F Figure 7 Main o r i g i n a r e a shown i n f i g u r e 6 a f t e r e x t e n s i v e c l e a n i n g by r e p l i c a t i o n . X300 ~ i F i g u r e 8. Longitudinal metallographic microsection through t h e main o r i g i n a r e a showing t h e d i s c o n t i n u i t y i n p r o f i l e . Arrows "R" and "F" i n d i c a t e t h e r a d i u s and f r a c t u r e p r o f i l e r e s p e c t i v e l y . 9x375 Kellers e t c h
I \ \
- 31 -
, APIJENDIX F Figure 9. Longitudinal section through.the fitting radius at a location diametrically opposite the fatigue origin. X10 unetched Figure 10. Longitudinal section through the main fatigue origin area. Arrow " 0 " indicates same area depicted by arrow " 0 " in figure 8. X10 nital etched