In-Flight Loss of Propeller Blade, Forced Landing, and Collision with Terrain, Atlantic Southeast Airlines, Inc., Flight 529, Embraer EMB-120RT, N256AS, Carrollton, Georgia, August 21, 1995

PB96-910406

NTSB/AAR-96/06

DCA95MA054

NATIONAL

TRANSPORTATION

SAFETY

BOARD

WASHINGTON,

D.C.

20594

AIRCRAFT ACCIDENT REPORT

IN-FLIGHT LOSS OF PROPELLER BLADE

FORCED LANDING, AND COLLISION WITH TERRAIN

ATLANTIC SOUTHEAST AIRLINES, INC., FLIGHT 529

EMBRAER

EMB-120RT,

N256AS

CARROLLTON,

GEORGIA

AUGUST 21, 1995

6609B

The National Transportation Safety Board is an independent Federal agency dedicated to

promoting aviation, railroad, highway, marine, pipeline, and hazardous materials safety.

Established in 1967, the agency is mandated by Congress through the Independent Safety

Board Act of 1974 to investigate transportation accidents, determine the probable causes of

the accidents, issue safety recommendations, study transportation safety issues, and evaluate

the safety effectiveness of government agencies involved in transportation. The Safety

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RE-51

490

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Road

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(703)487-4600

NTSB/AAR-96/06 PB96-910406

NATIONAL TRANSPORTATION

SAFETY BOARD

WASHINGTON, D.C. 20594

AIRCRAFT ACCIDENT REPORT

IN-FLIGHT LOSS OF PROPELLER BLADE

FORCED LANDING

AND COLLISION WITH TERRAIN

ATLANTIC SOUTHEAST AIRLINES, INC., FLIGHT 529

EMBRAER EMB-120RT, N256AS

CARROLLTON, GEORGIA

AUGUST 21, 1995

Adopted: November 26, 1996

Notation 6609B

Abstract:

This report explains the accident involving Atlantic Southeast Airlines flight

529, an EMB-120RT airplane, which experienced the loss of a propeller blade and

crashed during an emergency landing near Carrollton, Georgia, on August 21, 1995.

Safety issues in the report focused on manufacturer engineering practices, propeller

blade maintenance repair, propeller testing and inspection procedures, the relaying of

emergency information by air traffic controllers, crew resource management training, and

the design of crash axes carried in aircraft. Recommendations concerning these issues

were made to the Federal Aviation Administration.

this page intentionally left blank

iii

CONTENTS

EXECUTIVE SUMMARY .................................................................. v

1. FACTUAL INFORMATION

1.1 History of Flight ..................................................................................... 1

1.2 Injuries to Persons .................................................................................. 5

1.3 Damage to Airplane................................................................................ 5

1.4 Other Damage......................................................................................... 6

1.5 Personnel Information ............................................................................ 6

1.6 Airplane Information.............................................................................. 7

1.6.1 General.................................................................................................... 7

1.6.2 Weight and Balance................................................................................ 8

1.6.3 Propeller Design ..................................................................................... 8

1.6.4 Airplane and Propeller Design Requirements on Released Blades....... 10

1.6.5 Airworthiness Standards for Engines and Propellers............................. 12

1.7 Meteorological Information ................................................................... 13

1.8 Aids to Navigation.................................................................................. 15

1.9 Communications..................................................................................... 15

1.10 Airport Information ................................................................................ 16

1.11 Flight Recorders ..................................................................................... 16

1.12 Wreckage and Impact Information......................................................... 17

1.12.1 Fuselage .................................................................................................. 18

1.12.2 Wings...................................................................................................... 19

1.12.3 No. 1 (Left) Engine Nacelle ................................................................... 19

1.12.4 No. 2 (Right) Engine Nacelle................................................................. 20

1.13 Medical and Pathological Information................................................... 20

1.14 Fire.......................................................................................................... 21

1.15 Survival Aspects..................................................................................... 21

1.16 Tests and Research ................................................................................. 23

1.16.1 Laboratory Examination of the Fractured Propeller Blade.................... 23

1.16.2 Previous Failures of Similar Model Propellers...................................... 26

1.16.3 Blade Inspection and Repair - Actions Taken by

Hamilton Standard and the FAA........................................................... 28

1.16.4 The Failed Propeller, Information, and Service History........................ 36

1.16.5 Results of 14RF-9/EMB-120 Stress Survey .......................................... 39

1.17 Organizational and Management Information....................................... 44

1.17.1 Hamilton Standard Division, United Technologies Corporation .......... 44

1.17.2 Hamilton Standard Propeller Customer Service Center ........................ 45

1.17.2.1 Employee Training at Hamilton Standard Customer Service Center.... 45

1.17.3 Designated Engineering Representative ................................................ 46

1.17.4 FAA Certification Engineer ................................................................... 47

1.18 Additional Information........................................................................... 48

1.18.1 Safety Board Recommendations ............................................................ 48

1.18.2 Postaccident Hamilton Standard and FAA Actions............................... 50

2. ANALYSIS

2.1 General.................................................................................................... 53

2.2 Analysis of the Propeller Blade Failure ................................................. 53

2.2.1 The Accident Blade’s June 1994 Inspection, Repair,

and Return to Service ........................................................................... 54

2.2.1.1 Inappropriate Use of PS960A Blending Repair .................................... 54

2.2.1.2 Sanding (Blending) of the Accident Blade ............................................ 56

2.2.2 Adequacy of Hamilton Standard Procedures for Detecting Corrosion . 59

2.2.2.1 Borescope Inspection.............................................................................. 59

2.2.2.2 Technician Training and Supervision .................................................... 60

2.2.2.3 Adequacy of Improvements to Inspection Procedures .......................... 61

2.3 Effect of Blade Resonance ..................................................................... 62

2.4 Adequacy of Vibration Testing.............................................................. 65

2.5 Effect of Blade Failure and Analysis of Terminating Action................ 66

2.6 FAA Oversight........................................................................................ 68

2.6.1 Role of Designated Engineer Representative and FAA

Certifying Engineer ............................................................................... 68

2.7 Weather................................................................................................... 69

2.8 Air Traffic Control Services................................................................... 70

2.9 Survival Factors Aspects........................................................................ 72

2.9.1 Time Management During Emergencies................................................ 72

2.9.2 Crash Axes.............................................................................................. 74

3. CONCLUSIONS

3.1 Findings .................................................................................................. 75

3.2 Probable Cause ....................................................................................... 79

4. RECOMMENDATIONS ..................................................................... 80

5. APPENDIXES

Appendix A--Investigation and Hearing................................................ 83

Appendix B--Cockpit Voice Recorder Transcript ................................. 84

Appendix C--Fracture Summary............................................................ 110

Appendix D--EMB-120/14RF-9 Stress Resurvey

................................. 111

EXECUTIVE SUMMARY

On August 21, 1995, about 1253 eastern daylight time, an

Empresa Brasileira de Aeronautica S. A. (Embraer) EMB-120RT, N256AS,

airplane operated by Atlantic Southeast Airlines Inc., (ASA) as ASE flight

529, experienced the loss of a propeller blade from the left engine propeller

while climbing through 18,100 feet. The airplane then crashed during an

emergency landing near Carrollton, Georgia, about 31 minutes after

departing the Atlanta Hartsfield International Airport, Atlanta, Georgia.

The flight was a scheduled passenger flight from Atlanta to Gulfport,

Mississippi, carrying 26 passengers and a crew of 3, operating according to

instrument flight rules, under the provisions of Title 14 Code of Federal

Regulations Part 135. The flightcrew declared an emergency and initially

attempted to return to Atlanta. The flightcrew then advised that they were

unable to maintain altitude and were vectored by air traffic control toward

the West Georgia Regional Airport, Carrollton, Georgia, for an emergency

landing. The airplane continued its descent and was destroyed by ground

impact forces and postcrash fire. The captain and four passengers sustained

fatal injuries. Three other passengers died of injuries in the following 30

days. The first officer, the flight attendant, and 11 passengers sustained

serious injuries, and the remaining 8 passengers sustained minor injuries.

The National Transportation Safety Board determines that the

probable cause of this accident was the in-flight fatigue fracture and

separation of a propeller blade resulting in distortion of the left engine

nacelle, causing excessive drag, loss of wing lift, and reduced directional

control of the airplane. The fracture was caused by a fatigue crack from

multiple corrosion pits that were not discovered by Hamilton Standard

because of inadequate and ineffective corporate inspection and repair

techniques, training, documentation, and communications.

Contributing to the accident was Hamilton Standard’s and the

Federal Aviation Administration’s failure to require recurrent on-wing

ultrasonic inspections for the affected propellers.

Contributing to the severity of the accident was the overcast

cloud ceiling at the accident site.

Safety issues in the report focused on manufacturer engineering

practices, propeller blade maintenance repair, propeller testing and

inspection procedures, the relaying of emergency information by air traffic

controllers, crew resource management training, and the design of crash

axes carried in aircraft. Recommendations concerning these issues were

made to the Federal Aviation Administration.

NATIONAL TRANSPORTATION SAFETY BOARD

WASHINGTON, D.C. 20594

AIRCRAFT ACCIDENT REPORT

IN-FLIGHT LOSS OF PROPELLER BLADE

FORCED LANDING

AND COLLISION WITH TERRAIN

ATLANTIC SOUTHEAST AIRLINES, INC., FLIGHT 529

EMBRAER EMB-120RT, N256AS,

CARROLLTON, GEORGIA

AUGUST 21, 1995

1. FACTUAL INFORMATION

1.1 History of Flight

On August 21, 1995, about 1253

eastern daylight time, an Empresa

Brasileira de Aeronautica S.A. (Embraer) EMB-120RT, N256AS, airplane

operated by Atlantic Southeast Airlines Inc., (ASA

) as ASE

flight 529,

experienced the loss of a propeller blade from the left engine propeller while

climbing through 18,100 feet. The airplane then crashed during an emergency

landing near Carrollton, Georgia, about 31 minutes after departing the Atlanta

Hartsfield International Airport (ATL), Atlanta, Georgia. The flight was a

scheduled passenger flight from ATL to Gulfport, Mississippi (GPT), carrying 26

passengers and a crew of 3, operating according to instrument flight rules (IFR),

under the provisions of Title 14 Code of Federal Regulations (CFR) Part 135. The

flightcrew declared an emergency and initially attempted to return to Atlanta. The

flightcrew then advised air traffic control (ATC) that they were unable to maintain

altitude and were vectored toward the West Georgia Regional Airport (CTJ),

Carrollton, Georgia, for an emergency landing. The airplane continued its

All times are reported in eastern daylight time unless noted.

The Atlantic Southeast Airlines Inc., corporate logo and airplane paint scheme are

represented by the letters ASA.

The air traffic control system call sign for flights of Atlantic Southeast Airlines is ASE.

Because a code sharing agreement existed between ASA and Delta Air Lines, passenger

flight schedules also identified the airplane as Delta flight 7529.

descent, passed through some trees, and was destroyed by impact forces with the

ground and postcrash fire. The captain and four passengers sustained fatal

injuries. Three other passengers died of injuries in the following 30 days. The

first officer, the flight attendant, and 11 passengers sustained serious injuries,

and

the remaining 8 passengers sustained minor injuries.

On August 21, 1995, the accident flightcrew began a 2-day trip at

Macon, Georgia (MCN). They operated the accident airplane, N256AS, as flight

ASE 211 from MCN to ATL. A jump seat rider, an ASA captain, reported that the

flight was uneventful and that the crew appeared to be rested and in a relaxed

mood during the flight.

In ATL, the captain remained in the airplane on the ground to receive

the ATC clearance; the first officer deplaned and remained in the immediate area.

The accident flight, ASE 529, was cleared IFR from ATL to GPT via the Atlanta 4

departure and flight planned route at flight level 240.

The estimated flight time

was 1 hour and 26 minutes. The ASA EMB-120 load manifest prepared by the

first officer recorded 26 passengers, 3 crewmembers, 724 pounds of cargo, and

2,700 pounds of fuel for departure.

ASE 529 taxied from the ramp area at 1210 and was airborne at 1223.

At 1236, the first officer reported to the west departure sector of the Atlanta air

route traffic control center (Atlanta Center) that they were climbing past 13,000

feet. About 1242, following several intermediate climb clearances, the controller

issued a clearance to climb and maintain flight level 240, which the flightcrew

acknowledged.

The flight data recorder (FDR) and the cockpit voice recorder (CVR)

data

indicated that at 1243:25, while climbing through 18,100 feet at 160 knots

indicated airspeed (KIAS), several thuds could be heard from the cockpit, and the

torque on the left engine decreased to zero. The airplane then rolled to the left,

pitched down, and subsequently started to descend. Immediately thereafter, the

FDR shows numerous flight control inputs consistent with an attempt to

counteract the flightpath deviations; however, the airplane attitude decreased to

Section 1.2 contains more details regarding serious injuries.

Flight level 240 represents a barometric altimeter indication of 24,000 feet.

Appendix B contains the transcript of the CVR. All relevant ATC communications with

ASE 529 are contained in the transcript.

about 9 degrees nose low, and the airplane began a descent rate that progressed to

about 5,500 feet per minute (fpm). The captain said, “I can’t hold this thing,” then

“help me hold it.” At 1244:26, the first officer declared an emergency with

Atlanta Center and stated, “we’ve had an engine failure.” Atlanta Center cleared

ASE 529 direct to the Atlanta airport.

According to data from the FDR and CVR, airspeed and descent rate

changes continued and were accompanied by abrupt excursions in vertical and

lateral acceleration values. At 1245:46, the CVR revealed that the first officer

informed the flight attendant that they had experienced an engine failure, had

declared an emergency and were returning to ATL, and he told her to brief the

passengers. At 1246:13, the first officer stated, “we’re going to need to keep

descending, we need an airport quick and uh, roll the trucks and everything for

us.” The controller provided the flightcrew with heading information to CTJ. The

flightcrew applied various combinations of flight control inputs and power on the

right engine, partially stabilizing the airplane descent rate to between 1,000 and

2,000 fpm and the airspeed to between 153 and 175 knots indicated airspeed

(KIAS).

The Atlanta Center controller lost ASE 529’s transponder code from

radar when the airplane descended through 4,500 feet. About 1250, he instructed

the flightcrew to contact Atlanta approach control. The flightcrew contacted

Atlanta approach and requested the localizer frequency and vectors for the West

Georgia Regional Airport. The controller issued the localizer frequency. The

flightcrew acknowledged and then requested vectors for a visual approach. The

controller verified the altitude of the airplane and that the flight was in visual

flight rules (VFR) conditions and said, “fly heading zero four zero...airport’s at

your about 10 o’clock six miles....” At 1251:47, ASE 529 acknowledged, “zero

four zero ASE 529.” This transmission was the last one received by the approach

controller from the accident flight. After 1251:30, airspeed steadily decreased

from 168 KIAS to about 120 KIAS. FDR and CVR information indicated that the

landing gear and flaps remained retracted. CVR sounds indicated that the first

ground impact occurred about 1252:45.

In postcrash interviews, survivors indicated that during the climbout,

they heard a loud sound and felt the airplane shudder. They also indicated that

two or three blades from the left propeller were wedged against the front of the

Figure 1.--Propeller installation - left wing.

wing. The flight attendant said that she looked out the left side of the aircraft and

observed, “a mangled piece of machinery where the propeller and the front part of

the cowling was.” Other passengers observed the propeller displaced outboard

from its original position on the engine (see Figure 1). The flight attendant stated

that after the first officer notified her of the flight’s emergency return to ATL (at

1245:46), she prepared the cabin for an emergency landing and evacuation. She

stated that she had no further dialogue with the flightcrew.

Investigators found the left engine propeller assembly early in the

ground debris path. The propeller hub contained three complete blades and about

1 foot of the inboard end of the fourth blade protruding from the hub. The

remainder of the fourth blade was not at the accident site. (See Section 1.12 for

more wreckage information.)

The accident took place in daylight visual conditions. The crash site

was located at 33 degrees, 34’, 50.5” north latitude and 85 degrees, 12’, 51.2” west

longitude. A topographical map indicated that the elevation of the site was 1,100

feet above sea level.

1.2 Injuries to Persons

Injuries

Flightcrew Cabincrew Passengers Other Total

Fatal 1 0 7 0 8

Serious 1 1 11

0 13

Minor 0 0 8 0 8

None 0

0 0 0 0

Total 2 1 26 0 29

1.3 Damage to Airplane

The airplane was destroyed by the impact and postcrash fire. Its

estimated value was $5,000,000.

One passenger died 4 months after the accident as a result of her injuries. She sustained

third-degree burns over 50 percent of her body, as well as inhalation injuries. In accordance with

49 CFR 830.2, which defines “fatal injury” as any injury that results in death within 30 days of

the accident, her injuries were classified as “serious.”

1.4 Other Damage

The crash site was located on 20 acres of unimproved farmland with

trees, adjacent to an open field. There was environmental damage from airplane

fuel and fire-fighting efforts along the wreckage path, and immediately adjacent to

the wreckage.

1.5 Personnel Information

The captain, age 45, held an airline transport pilot certificate for

airplane multiengine land, type rated in the EMB-120, with commercial privileges

for airplane single-engine land. He held a flight instructor certificate with ratings

for airplane, instrument, and multiengine. His most recent Federal Aviation

Administration (FAA) first-class medical certificate was issued on April 3, 1995,

with the limitation: “Holder shall wear correcting lenses for near vision while

exercising the privileges of his airman certificate.” The captain’s overnight bag,

found in the wreckage, contained an empty eyeglasses case.

The captain was employed by ASA in March 1988. Company records

indicate that at the time of the accident, he had accumulated 9,876.13 total hours

of flying experience, with 7,374.68 hours in the EMB-120 of which 2,186.94

hours was pilot in command. His last proficiency check was on March 3, 1995,

and his most recent training, on August 7, 1995, was Line Oriented Flight

Training (LOFT).

The first officer,

age 28, held a commercial pilot certificate with

ratings for airplane, single-engine land, airplane multiengine land, and instrument-

airplane. He held a flight instructor certificate with ratings for airplane,

multiengine, and instrument. His most recent FAA first-class medical certificate

was issued on June 15, 1995, without limitations.

The first officer was employed by ASA in April 1995. Company

records indicate that at the time of the accident, he had accumulated 1,193 total

hours of flying experience, with 363 hours in the EMB-120. He received his ASA

first officer training in April 1995, and completed his initial operating experience

on May 4, 1995.

Because of his severe injuries that included burns and inhalation damage, Safety Board

investigators were unable to interview the first officer.

The flight attendant, age 37, was employed by ASA on February 8,

1993. She completed her initial training, which included emergency procedures

training, on February 23, 1993. She had no prior experience as a flight attendant.

Her most recent recurrent training on the EMB-120 was on January 26, 1995.

Activities of the crew in the days before the accident were routine and

unremarkable. They appeared to have received normal rest.

1.6 Airplane Information

1.6.1 General

The airplane, N256AS, was an Embraer EMB-120RT “Brasilia,”

serial number 120122, manufactured and certificated in Brazil. The airplane was

certificated in the United States in accordance with a bilateral airworthiness

agreement between the FAA and the Brazilian certification authorities. The

airplane was delivered to ASA on March 3, 1989.

Prior to the day of the accident, the airplane had accumulated

17,151.3 flight hours and 18,171 flight cycles. Maintenance records indicate that

maintenance inspections were accomplished in accordance with ASA’s Standard

Practice No. 624, Airplane Maintenance Program, an FAA-approved maintenance

plan.

The airplane had been assigned to ASA’s Dallas-Fort Worth, Texas,

hub until 1 week prior to the accident when it was transferred to ASA’s Atlanta

hub in preparation for a “C” check (required at 3,300-hour intervals). The

inspection was scheduled to take place the coming week at the ASA maintenance

facility at MCN but the accident intervened. Safety Board investigators reviewed

the airplane records at MCN and found no remarkable discrepancies or minimum

equipment list (MEL) items carried forward in the records.

As part of the ASA maintenance program, a line check

is required at

75-hour intervals. A line check was accomplished by employees of ASA’s MCN

maintenance facility on the night of August 20, 1995, during an overnight stop.

According to ASA maintenance work cards, the line check included a specific

An ASA line check for the EMB-120 airplane consists of detailed visual inspections in

14 areas to detect leaks, damage, and ensure the continuing airworthiness of all systems.

visual inspection of the left and right propellers for any evidence of oil leaks or

damage; and none were noted. Maintenance records indicate that a maintenance

daily inspection

was performed at MCN on August 21, 1995, the date of the

accident, by ASA line mechanics prior to the first flight of the day. Maintenance

personnel indicated that a flightcrew walk around inspection was also

accomplished by the first officer prior to departing MCN. Neither of these

inspections noted anything remarkable.

1.6.2 Weight and Balance

ASA dispatch records indicate that the takeoff weight of N256AS at

ATL was 24,237 pounds. The maximum takeoff weight, as stated in the airplane

flight manual (AFM), is 25,353 pounds. The planned landing weight was 22,637

pounds; the AFM maximum landing weight is 24,802. The takeoff percent mean

aerodynamic chord (MAC) was 28.65; the AFM forward and aft center of gravity

limits are 21.0 and 42.0 percent MAC respectively. The airplane was within its

prescribed weight and center of gravity limitations at takeoff and at the time of the

accident.

1.6.3 Propeller Design

Hamilton Standard manufactures a family of composite propeller

blades, including the 14RF (the accident propeller blade), 14SF, and 6/5500/F,

that are intended for use on turbopropeller commuter airplanes. The solid, forged

7075-T73 aluminum alloy spar is the main load-carrying member. The airfoil

shape of the blade is formed by glass fiber-filled epoxy and foam adhesively

bonded to the spar (see Figure 2). A conical hole (taper bore) is bored in the

center of the spar from the inboard end to blade station 21,

for weight reduction

and balance weight installation. The taper bore on the 14RF blade can be one of

two different shaped designs: straight taper bore known as the “M” style

) and a

An ASA maintenance daily inspection is performed prior to the first flight of the day. It

consists of external and internal visual inspections, checks of system operating pressures and

fluid levels, and an operational check of radio and navigational equipment.

Blade stations on the 14RF model propeller are measured in inches from a reference

point 3.427 inches inboard of the blade pin platform on the inboard end of the blade.

Produced from February 1986 to February 1987 through serial number 85344.

63.00 STA

60.00 STA

LEADING EDGE

NICKEL SHEATH

48.00 STA

36.00 STA

EROSION STRIP

DE-ICE HEATER

24.00 STA —

16.60 STA —

12.00 STA

;

00.00 STA

HAMILTON STANDARD

14RF-9

PROPELLER BLADE

UNTWISTED PLANFORM

WITH SECTION THRU

SHANK & TAPERBORE

OF THE ALUMINUM SPAR

ALUMINUM SPAR

SEPARATION

TAPER BORE

Figure 2.--Illustration of a

14RF-9

propeller blade.

bellmouth shape (known as the “N” style). Also, during very early production, the

taper bore was shotpeened.

However, early in the production run, Hamilton

Standard reviewed the production process, deemed shotpeening unnecessary, and,

with FAA approval, it was discontinued. The accident blade was originally an N

style blade, but it was rebalanced on customer request for ASA fleet

standardization to the M style and reidentified as an M blade. The accident blade

taper bore was not shotpeened during production.

The taper bore provides space for a measured amount of lead wool to

be inserted for blade balancing. Until April 1994, a cork was used to retain the

lead wool in the taper bore; however, it was later eliminated

because it was found

to be a source of chlorine (and potential corrosion) in the taper bore, as well as

unnecessary to retain the lead wool. The model 14RF, 14SF and 6/5500/F

propellers, which vary in length from 10.5 to 13 feet, operate at a maximum of

1,200 to 1,384 rpm. As of July 1996, the 14RF, 14SF and 6/5500/F propeller

assemblies were installed on 9 types of commuter aircraft, operated by 143

operators, on approximately 1,300 aircraft, for an industry total of about 15,000

blades worldwide.

According to information provided to the Safety Board in February,

1995, Hamilton Standard statistical data from field service experience indicate that

blades without shotpeened taper bores are susceptible to earlier corrosion and

cracking.

1.6.4 Airplane and Propeller Design Requirements on Released Blades

When the EMB-120 was certificated in the United States, the effect

on safe flight of a failed or released propeller blade was addressed in FAR

25.571(e)(2), “Damage tolerance (discrete source) evaluation,” which, at that time,

required that the airplane be capable of successfully completing a flight during

which likely structural damage occurs as a result of a propeller blade impact.

Shotpeening is a metallurgical surface treatment to improve resistance to cracking. The

surface to be treated is bombarded with air-propelled glass beads or steel shot. Only the first 431

production blades were shotpeened.

Use of cork was discontinued in the manufacturing process between April 1994 and

November 1995. Existing corks have been removed from all model 14RF blades, and will be

removed from other models pursuant to PS960A and later AD 95-05-03, which sets forth end

dates for each propeller model.

Embraer petitioned the FAA on January 11, 1983, to permit type certification of

the EMB-120 without compliance with that requirement. On March 22, 1983, the

FAA exempted the EMB-120 from compliance with FAR 25.571(e)(2) through

Grant of Exemption No. 3722, which also contained the statement that, “all

practical precautions must be taken in the design of the airplane taking into

account the design features of the propeller and its control system to reduce the

hazard which might arise from failure of a propeller hub or blade.”

In 1990, the (e)(2) requirement was eliminated from FAR 25.571.

The FAA noted at that time (in the preamble to this and other regulatory changes),

that service experience had shown compliance to be “impossible,” and that “[a]s a

result of the granting of exemptions for good cause, no manufacturer has, in fact,

been required to show compliance with the current requirement.” At the same

time, however, the FAA promulgated FAR 25.905(d), which requires that “design

precautions must be taken to minimize hazards to the airplane from a failed or

released blade, including damage to structure and vital systems due to impact of a

failed or released blade, and from the unbalance created by such failure.” (55 Fed.

Reg. 29756 at 29772, 29766, July 20, 1990.)

According to Embraer records, after initial certification of the

airplane, Embraer evaluated the effects on the wing, the nacelle, and the

empennage from a blade tip loss, a mid-blade loss, and a full blade loss.

Embraer’s analysis indicated that the nacelle would not withstand the loss of a

mid-blade or full blade segment.

After the accident, the FAA indicated that records from the original

certification of the 14RF propeller indicate that Hamilton Standard demonstrated

compliance, through testing, with the requirement of FAR 35.15 that the propeller,

“not have design features that experience has shown to be hazardous or

unreliable.” The FAA further stated that Hamilton Standard designed, tested and

demonstrated the 14RF-9 propeller blade to meet the FAR 35 requirements and it

was approved as having an unlimited life when maintained in accordance with

FAA-accepted Hamilton Standard maintenance instructions.

Embraer Report No. 120-DE-180, “Effect Analysis on Propeller Failures,” dated

September 3, 1984.

1.6.5 Airworthiness Standards for Engines and Propellers

To comply with the airworthiness requirements, the propeller

manufacturer must also consider during design and must subsequently

demonstrate the vibration characteristics of the propeller assembly to ensure that

resonant frequencies

that can produce critical vibration stresses do not occur

within the normal operating range of use. The applicable regulations are 14 CFR

23.907,

25.907,

35.37,

and 35.39.

Advisory Circular

(AC) 20-66 provided

The resonant frequency of any vibration is the naturally occurring frequency at which

the blade will vibrate when excited. To avoid excessive vibration and overstressing of the

propeller, propeller design practice requires that the propeller spend only a minimal amount of

time in an rpm range that corresponds to a resonant frequency.

Each propeller with metal blades or highly stressed metal components must be shown to

have vibration stresses, in normal operating conditions, that do not exceed values that have been

shown by the propeller manufacturer to be safe for continuous operation. This must be shown

by: Measurement of stresses through direct testing of the propeller; comparison with similar

installations for which these measurements have been made; or any other acceptable test method

or service experience that proves the safety of the installation. Proof of safe vibration

characteristics for any type of propeller, except for conventional, fixed pitch, wood propellers,

must be shown where necessary.

The magnitude of the propeller blade vibration stresses under any normal condition of

operation must be determined by actual measurement or by comparison with similar installations

from which these measurements have been made. The determined vibration stresses may not

exceed values that have been shown to be safe for continuous operation.

A fatigue evaluation must be made, and the fatigue limits must be determined for each

metallic hub and blade and each primary load-carrying metal component of nonmetallic blades.

The fatigue evaluation must include consideration of all reasonably foreseeable vibration load

patterns. The fatigue limits must account for the permissible service deterioration, such as nicks,

grooves, galling, bearing wear, and variations in material properties.

For variable-pitch propellers. Compliance with this paragraph must be shown for a

propeller of the greatest diameter for which certification is requested. Each variable-pitch

propeller (the pitch setting can be changed by the flightcrew or by automatic means while the

propeller is rotating) must be subjected to one of the following tests: A 100-hour test on a

representative engine with the same or higher power and rotational speed and the same or more

severe vibration characteristics as the engine with which the propeller is to be used. Each test

must be made at the maximum continuous rotational speed and power rating of the propeller. If

a takeoff rating greater than the maximum continuous rating is to be established, an additional

10-hour block test must be made at the maximum power and rotational speed for the takeoff

rating. Operation of the propeller throughout the engine endurance tests is prescribed in Part 33

of this subchapter.

the propeller manufacturer an acceptable means of compliance with the CFRs

relating to airplane propeller vibration.

In Chapter Two, Vibration Measurement Program, it is recommended

that for multiengine installations:

Propeller diameters to be used are tested in at least two percent or

two-inch intervals throughout the diameter range to be approved

and should include the maximum diameter and the minimum

diameter, including cutoff repair limit.

AC 20-66 also states:

For installations where the propeller diameter is greater than 13

feet or the engine nacelles are toed in or toed out, propeller

vibration testing include complete flight and ground crosswind

tests. Flight tests includes the effects of yaw, maximum and

minimum aircraft gross weight at maximum and minimum

airspeeds, flap settings during takeoff and landings, propeller

reversing, and any other condition that would create an

aerodynamic excitation of the propellers. On the ground, the

aircraft is headed at different angles to the prevailing wind to

determine the effects of crosswind excitation. Wind velocities

typical of conditions to be encountered in service are included.

1.7 Meteorological Information

The West Georgia Regional Airport (CTJ) at Carrollton, Georgia, is

about 4 miles northeast of the crash site. The airport authority owns and operates

an Automated Weather Observing System-3 (AWOS-3). The CTJ AWOS-3, and

similar independent systems at other airports that do not serve air carriers, are not

connected through long-line transmission to the National Weather Service (NWS)

or the FAA weather communication networks. The observations are available to

airport users on a dedicated radio frequency. The CTJ AWOS-3 observation just

after the accident was reported as follows:

An AC is an FAA document that sets forth an acceptable means to comply with

provisions of Federal Aviation Regulations (FAR). An AC is intended for guidance purposes

only and is not mandatory or regulatory in nature.

Type--AWOS-3; time--1301; clouds--800 feet overcast; visibility--

10 miles; temperature--76 degrees F; dew point--75 degrees F;

wind--150 degrees at 6 knots; altimeter--30.08 inches Hg;

Anniston Airport (ANB), Alabama is about 32 miles west of the crash

site. The reported ANB aviation weather observation just before the accident was

as follows:

Type--Record; time--1252; clouds--estimated ceiling 1,500 feet

broken; visibility--5 miles; weather--haze; temperature 87 degrees

F; dew point--73 degrees F; wind--050 degrees at 5 knots;

altimeter--30.02 inches Hg.

The departure airport, ATL, is about 40 miles east of the crash site.

The reported ATL aviation weather observation just before the accident was as

follows:

Type--Record special; time--1246; clouds--200 feet scattered

measured ceiling 1,600 feet broken 3,400 feet overcast; visibility--

2 miles; weather--light rain fog, temperature 73 degrees F; dew

point--73 degrees F; wind--140 degrees at 3 knots; altimeter--

30.08 inches Hg; remarks--surface visibility 5 miles.

The CVR transcript at 1250:15 (2 minutes and 30 seconds before

impact) contained a captain’s statement that, “we can get in on a visual.” The

FDR altitude at that time was about 3,760 feet. The CVR transcript at 1251:05 (1

minute and 40 seconds before impact) contained a captain’s statement, “we can get

in on a visual, just give us vectors.” The FDR altitude at that time was about

2,450 feet. The ATC and CVR transcripts indicate that the first officer reported at

1251:33 (1 minute and 12 seconds before impact) “out of nineteen hundred (feet)

at this time” and the captain added “we’re below the clouds, tell ‘m.” The first

officer then transmitted, “’K we’re uh, VFR at this time, give us a vector to the

airport.”

A helicopter pilot, who arrived at the accident site about 1400,

estimated scattered clouds at about 1,500 feet and a broken ceiling at around 2,500

feet. He estimated the visibility at 3.5 miles in haze.

1.8 Aids to Navigation

There were no reported or known difficulties with the navigational

aids.

1.9 Communications

There were no reported or known communications equipment

difficulties.

At the time of the propeller blade separation, ASE 529 was

communicating with an Atlanta Center air traffic controller. Although the base of

the Atlanta Center controller’s airspace is 11,000 feet, the center controller

continued to direct the airplane for about 7 minutes after the blade separation. At

that point (1250:45), with the airplane at about 4,500 feet in altitude, changeover

to Atlanta approach took place. Recorded radar information at that time indicated

that the airplane was about 7 miles from CTJ. The Atlanta approach controller

issued a vector toward the CTJ ILS localizer at 1250:49. Later the controller

provided the localizer frequency; however, neither the AWOS frequency nor the

CTJ weather conditions were provided.

Atlanta approach is responsible for flights inbound or outbound from

ATL and all airports within an approximate 40-mile radius, which includes CTJ.

However, the ATL approach control facility’s access to the CTJ AWOS weather

information is limited to commercial telephone sources. The Georgia Department

of Transportation, which (as the operator of the airport) would be responsible for

the costs of disseminating AWOS information via private communications

networks directly to ATC, determined that the amount of air traffic at CTJ did not

justify the cost of acquiring this service. This is because flightcrews destined for

the smaller airports receive their AWOS weather information on the airport

discrete AWOS frequency.

The closest weather report immediately available to the approach

controller was the ATL Airport observation, the flight’s departure point. No

controller was assigned to the “assist” position. Although the manager and

supervisor were nearby, they became occupied with coordinating and monitoring

activities supporting the flight and did not attempt to retrieve the CTJ AWOS

weather information by telephone. During the 90 seconds that the approach

controller was in radio communication with the flight, the controller issued a

vector toward the runway, stated the localizer frequency, confirmed the flight was

in visual conditions, and issued a vector for the visual approach.

FAA ATC procedures

state, in part, “If you are in communication

with an aircraft in distress, handle the emergency and coordinate and direct the

activities of assisting facilities. Transfer this responsibility to another facility only

when you feel better handling of the emergency will result.”

Following their declaration of an emergency with Atlanta Center, at

1246:13 the flightcrew indicated their need to land as soon as possible and

requested, “roll the trucks and everything for us.” The controller then advised the

flightcrew that CTJ was the closest airport and directed the aircraft to CTJ.

However, the request for emergency vehicles was not passed to the fire department

serving CTJ, (the Carroll County Fire Department) or to the Atlanta approach

controller. Following the accident, Atlanta approach did call the Carroll County

Sheriff’s Office and was informed that a private citizen had already reported the

airplane crash near CTJ.

1.10 Airport Information

CTJ has one asphalt surface runway, 5,001 feet by 100 feet, oriented

340/160 degrees, and the field elevation is 1,160 feet. There are two instrument

approaches, an instrument landing system (ILS) localizer only (LOC) RWY 34

and a nondirectional beacon (NDB) RWY 34. Atlanta approach control is the

feeder ATC agency on sector frequency 121.0 megahertz (MHz). Weather at the

airport is available directly through an AWOS-3 reporting system on 118.175

MHz. The airplane crashed about 4 miles from the airport.

1.11 Flight Recorders

The airplane was equipped with an operating cockpit voice recorder

(CVR) and flight data recorder (FDR). They were recovered from their installed

positions in the aft portion of the airframe and appeared in good condition with

only minor sooting on the cases. The CVR was a Fairchild Model A100A, S/N

57597. The recording was good and showed no evidence of loss of quality as a

result of the crash.

FAA Order 7110.65, Chapter 10, “Emergencies,” Section 1, “General,” paragraph 10-1-

4, “Responsibility,” applies.

The FDR was a Fairchild Digital Model F-800, S/N 04856, with 28

parameters of data. The recording was of good quality; however, the parameter

for rudder pedal position indicated only small changes that did not approach

normal travel. Postaccident investigation of the airframe wreckage revealed that

the rudder pedal potentiometer coupler was not securely connected to the shaft of

the rudder pedal potentiometer.

ASA maintenance records indicate that the rudder pedal

potentiometer was installed on the accident airplane in November 1990. The most

recent calibration check was performed in June 1994. At that time, no

discrepancies were noted during a 3-point calibration check (neutral, full left, and

full right). On June 27, 1996, the Safety Board issued two safety

recommendations to address this issue (see Section 1.18.1).

1.12 Wreckage and Impact Information

The main wreckage area consisted of the cockpit, fuselage, right wing

and engine, and the empennage. Portions of two of the right engine’s propeller

blades remained attached to the propeller hub and engine. The remaining two

blades of the right engine propeller assembly were located nearby. An area of the

grass leading up to and surrounding the main wreckage was burned out to a radius

of about 30 feet.

The airplane came to rest at the northwest end of an 850 foot

wreckage trail that was aligned on a heading of about 330 degrees magnetic.

Numerous trees were sheared off prior to ground contact, consistent with a descent

angle of about 20 degrees, and an increasing left-wing-down attitude of 15 to 40

degrees. Impact with the trees extended for about 360 feet, and, following the last

tree impact, a debris path continued for 490 feet through an open field on slightly

upsloping terrain to the main wreckage.

Prominent ground scars were observed at the beginning of the debris

field (about 40 feet from the last tree impact) that were consistent with the

dimensional measurement of the left wing to the fuselage. The first scars

contained several pieces of the left wing. Ground scars were consistent with

separation of the left wing at its root. Debris from the airplane was scattered along

the wreckage path in the field. The left engine’s propeller and reduction gear box

(RGB) assembly were located approximately 160 feet past the tree line. The

propeller hub and blade assembly contained three complete propeller blades with

the inboard piece of a fourth blade protruding about 1 foot from the hub.

The Safety Board’s Airplane Performance Group used its

WINDFALL computer program to calculate the trajectory of the missing blade

piece. The group devised a search area and alerted the local residents and

authorities about the missing piece. Three weeks after the accident, the outboard

piece of the blade was discovered by a farmer. It had been well hidden in some

tall grass within about 100 yards of the primary search area.

The fractured blade sections were sent to the Safety Board’s Materials

Laboratory for detailed examination (see Section 1.16.1 for details).

1.12.1 Fuselage

The aft portion of the fuselage had separated from the forward portion

in two places, near the trailing edge of the wing and also just behind the cockpit.

The forward fuselage section (including the cockpit) was upright. The aft portion

of the fuselage was resting on the right side and was supported by the right

horizontal stabilizer. The vertical stabilizer was intact and essentially undamaged.

Most of the passenger cabin that was not resting on the ground was destroyed by

fire.

The right side of the forward fuselage from the radome rearward to

the cockpit had very little damage. The left side of the forward fuselage below the

cockpit window from the radome to just forward of the passenger/crew entry door

was crushed in, aft, and up to the left side of the nose landing gear wheel well.

Inward deformation was less severe near the aft portion of the crushed area. The

external fuselage skin forward of the passenger/crew entry door was undamaged

by fire except for an area of sooting aft and above the captain’s side window.

Fire had destroyed the left side of the fuselage aft of the

passenger/crew entry door. The fire damage extended to just forward of the cargo

door and the entire right side of the fuselage from the leading edge of the wing to

two seat rows forward of the cargo section. The upper portion of the right

fuselage forward of the leading edge of the wing to the cockpit had also been

destroyed by fire.

1.12.2 Wings

The major portion of the left wing, with the nacelle and engine

partially attached, came to rest along the wreckage path about 125 feet in front of

the cockpit. The inboard portion of the left wing leading edge, from the fuselage

to the left engine nacelle, was intact. The leading edge outboard of the left engine

nacelle was recovered from the debris field but was broken into several pieces.

There were no cuts or gouges in the leading edge. The inboard and nacelle flaps

and the inboard flap track for the outboard flap were attached. Damage to the flap

tracks was consistent with the flaps being in the retracted position.

The entire right wing remained intact and attached to the fuselage.

The inboard section of wing between the engine and the fuselage was destroyed by

fire. There was no fire damage to the wing outboard of the engine. All flap

segments appeared to be in the retracted position.

1.12.3 No. 1 (Left) Engine Nacelle

The outboard member of the front frame of the No. 1 left engine

nacelle was deformed aft approximately 90 degrees and was twisted outboard

slightly. There was also a semicircular flattened area in the middle of the outboard

member of the front frame. The axis of the flattened area was oriented upward

approximately 20 degrees from the horizontal. The forward inboard engine mount

bolt had sheared in an upward and slightly outboard direction. The corresponding

metal surface area of the attachment fitting was smeared in the same direction.

The engine air inlet fairing and the forward portion of the forward

cowling remained attached to the propeller/RGB assembly, but they were

deformed outboard. Both steel tubes connected to the forward and aft engine

mounts were found separated from the terminal ends. The inboard tube was bent

slightly; the outboard tube was not bent.

Five of the six hinges that secure the inboard and outboard forward

cowling doors were attached, but they were bent in a direction consistent with up

and aft movement of the cowling doors. The area underneath several of the hinges

was damaged consistent with overtravel of the hinges. The forward, inboard hinge

had separated, and the area of the inboard door where the hinge was attached was

torn. The forward edge of both forward cowling doors was bent upward.

The forward inboard engine/RGB mount bolt, and forward, outboard

engine/RGB mount, upper and lower rod ends of the inboard and outboard torque

mount assemblies were removed and submitted to the Safety Board’s Materials

Laboratory for examination of all fracture surfaces. That examination revealed no

indications of fatigue or other preexisting defects. The inboard engine/RGB

mount and the outboard engine mount bolt were intact and remained attached to

the engine and the nacelle structure, respectively. No deformation was noted on

the inboard engine mount. The forward, outboard engine/RGB mount was

deformed aft near the fracture location. No definitive failure directions were

obtained from the upper rod ends, which had fractured near the first screw thread.

Examination of the fracture surfaces of the lower rod ends revealed characteristics

consistent with the fracture propagating inboard to outboard.

1.12.4 No. 2 (Right) Engine Nacelle

The No. 2 engine and RGB remained mounted to the wing. Although

a fire consumed the adjacent inboard wing-to-fuselage area, damage to the No. 2

engine nacelle was not remarkable. All cowlings and fairings were found in place

with little evidence of fire or soot.

1.13 Medical and Pathological Information

The Carroll County Medical Examiner determined that the seven

fatally injured passengers succumbed to thermal burns and smoke inhalation. The

cause of death for the captain was also reported to be thermal burns and smoke

inhalation. However, in his report, the Medical Examiner indicated that blunt

force trauma injuries to the face and head were “other significant conditions.”

This is consistent with impact-related damage on the forward left side of the

fuselage. The first officer survived with burns over 80 percent of his body.

Physicians indicated that as a result of his injuries, he would require extensive

therapy.

Urine samples obtained post-mortem from the captain, and blood and

urine samples obtained from the first officer after the accident, tested negative for

alcohol and other drugs of abuse.

1.14 Fire

Based on ASA flight dispatch records, investigators estimated that

about 350 gallons of fuel were on board at the time of the accident. Per normal

operating practice, the fuel would have been equally distributed between the left

and right side tanks. The two tanks in the left wing separated early during the

impact sequence, and there was evidence of fuel spilled on vegetation along the

wreckage path. The inboard tank in the right wing was found burned at the

accident site, but the outboard tank was intact. Passengers did not observe fire

until after the airplane came to a complete stop. They said that there was a period

of about 1 minute before the outbreak of fire. The passengers described black

smoke and flame consistent with what would be expected of a fuel-fed fire.

Passengers reported that the fire was immediately preceded by cracking sounds

and sparks from wires and cables and that the fire started in small patches and

spread quickly, fully engulfing the area aft of the cockpit entrance door.

Some passengers related that they found portions of their clothing

saturated with fuel, and one passenger saw “a couple of people on fire.” The flight

attendant and several passengers said that they had to run through flames to escape

from the cabin wreckage.

The flight attendant received second degree burns to her ankles and

legs. She was wearing a skirt, white blouse, hosiery, and an airline uniform vest.

1.15 Survival Aspects

The CVR revealed that the flightcrew advised the flight attendant of

the planned emergency return to ATL about 7 minutes prior to impact. There were

no further communications from the flightcrew to the flight attendant. The flight

attendant stated that while preparing the passengers for the emergency landing,

she saw tree tops, immediately returned to her seat, and shouted commands to

brace for landing.

According to passengers, immediately following the loss of the

propeller blade, the flight attendant checked with each passenger to make sure that

they understood how to assume the brace position, and she yelled instructions to

the passengers until the time of impact. Despite being seriously injured, she

continued to assist passengers after the accident by moving them away from the

airplane. She also extinguished flames on at least one passenger who was on fire.

The postcrash fire destroyed the passenger cabin. According to the

surviving passengers, the cabin breakup started at the initial ground impact.

Passengers stated that overhead storage bins in the cabin dislodged during the

initial ground impact and that passenger seat structures separated and/or became

deformed. According to one passenger, as the fuselage slid on its left side, several

large holes were created that allowed enough daylight to appear in the cabin that

provided the flight attendant and passengers visual escape cues. None of the

survivors reported escaping from the cabin through the main entrance door, the

overwing emergency exits, or the cabin emergency exit. They escaped through the

holes in the fuselage, which were immediately behind the cockpit and aft of the

wing. Passengers who were unable to escape from the wreckage succumbed to

smoke inhalation.

Shortly after the airplane came to rest, the first officer attempted

unsuccessfully to open the right side cockpit window, which was damaged during

the impact. Thereafter, he reached behind his crew seat and retrieved a small ax

with a wooden handle. He subsequently attempted to chop a hole in the side

window but was only successful in chopping a hole approximately 4 inches in

diameter in the center of the window through which he handed the small ax to a

passenger. The passenger attempted unsuccessfully to use the ax to extricate him

from the cockpit.

When a Carroll County Sheriff’s deputy arrived at the scene within

about 5 minutes, he saw a passenger striking the first officer’s side window with

the small ax,

which was aboard the airplane as FAA-required emergency

equipment. The wooden handle separated from the ax head early in the rescue

effort. About 2 minutes after the ax handle broke, the local fire department arrived

and tried, unsuccessfully to break the window using full size axes. The fire

department applied water to the cockpit side window. The deputy reported that

during the time of the rescue, a continuous roaring sound emanated from an area

behind the cockpit in which there was intense fire. In the following several

minutes, the fire aft of the cockpit was controlled sufficiently to allow firefighters

to enter the cabin and break through the cockpit door to rescue the first officer.

The Sheriff’s deputy did not observe any signs of life from the captain during the

rescue sequence.

The ax had a short wooden handle about 14 inches long and resembled a hatchet. It had

a single blade with a nail puller notch, and the opposite end of the blade had a shape that was

similar to a hammerhead.

Postaccident inspection of the cockpit area indicated that movement

of the right and left cockpit sliding windows was restricted by airframe damage

consistent with impact and deformation of the windows’ slide tracks. The first

officer’s cockpit sliding window was found to have jammed in its track in the

closed position. Investigators were able to open the sliding windows with the aid

of pry bars (tools not normally available to flightcrews ).

The flightcrew oxygen walkaround cylinder and smoke masks were

found stored, respectively, on the left and right sides of the cockpit. They did not

appear to have been used. Protective breathing equipment (PBE) required in 14

CFR Part 121 airplanes was not carried (nor was it required to be) because the

airplane was operated under 14 CFR Part 135.

1.16 Tests and Research

1.16.1 Laboratory Examination of the Fractured Propeller Blade

The inboard piece of the fractured blade, serial number 861398, was

retained in the left engine propeller hub, which was recovered at the accident

scene on August 21, 1995. The outboard piece of the blade was recovered on

September 15, 1995, after it was discovered by a farmer on property about 35

miles west of the accident site adjacent to an area that had previously been

searched by helicopter. Both portions of the blade were examined at the Safety

Board’s Materials Laboratory. The blade spar

was separated at blade station 16.6

(about 13.2 inches outboard of the blade pin platform). Initial visual examination

revealed that a portion of the spar fracture was on a flat transverse plane and

contained crack arrest positions, typical of fatigue cracking (see Figure 3). The

fatigue cracking appeared to initiate from at least two adjacent locations on the

taper bore surface. Below the taper bore surface, the individual cracks joined to

form a single crack that propagated toward the face side

of the blade and

progressed circumferentially around both sides of the taper bore. The extent of the

fatigue cracking progressed through about 75 percent of the spar cross section.

The fracture surface in areas beyond the terminus of the fatigue region contained

rough features with a matte appearance, typical of an overstress separation area.

The main load-carrying member of the blade.

The face side of the blade is aerodynamically similar to the bottom side of a wing.

Figure 3.--Photo of the blade fracture surface.

The origin areas on both the inboard and outboard faces of the

fracture were examined with a scanning electron microscope (SEM) before the

faces of the fracture were cleaned. The fracture surface near the origin area on

both faces of the fracture contained a layer of heavy oxide deposits that had a

mud-cracked appearance. These deposits extended to a maximum depth of 0.049

inch from the taper bore surface. Their maximum circumferential width was 0.130

inch, based on the examination of the damaged outboard fracture face. X-ray

energy dispersive spectroscopy (EDS) of the deposits of both faces generated

similar spectra with a major peak for aluminum, a substantial peak for oxygen, and

a minor peak for zinc.

EDS of the deposit area on the inboard fracture face also

revealed the presence of chlorine.

After the fracture surface had been cleaned, additional SEM

examination revealed that the fatigue cracking initiated from several corrosion pits

in a line of pits in the taper bore that extended over a distance of about 0.070 inch.

The maximum depth of the corrosion pitting at the fatigue origin area was

measured as slightly less than 0.006 inch below the taper bore surface.

The taper bore surface, including the area adjacent to the fatigue

initiation area, contained a series of nearly circumferential sanding marks. The

marks extended over about 180 degrees of the circumference of the taper bore and

to a maximum distance of about 1.5 inches inboard of the fracture surface.

Outboard of the fracture, the sanding marks extended about 2 inches from the

fracture surface.

The investigation revealed that sanding rework of the area had been

accomplished using the blending repair procedures contained in PS960A.

The

procedures required that the surface finish of the blended area should be restored

to the original surface finish. Postaccident surface profilometer measurements

conducted on the taper bore sanding marks indicated that the surface finish was

much rougher than the manufactured surface not disturbed by the rework

process.

Measurements also indicated that the nondisturbed surface met the

manufacturing specifications.

Zinc is an alloying element in the 7075 aluminum alloy specified for the blade spar.

PS960A is described in paragraph 1.16.3.

The surface roughness in the blended area was measured as Ra 125, whereas the surface

finish requirement of PS960A is 63 RMS, which converts to approximately Ra 50. (“Ra”

denotes arithmetically averaged roughness.)

The taper bore was measured to determine the minimum thickness of

the spar between the taper bore hole and the spar’s face side and was found to be

within the requirements of the manufacturing specifications. Measurements also

indicated that about 0.002 inch of material appeared to have been removed from

the taper bore surface during the sanding process.

Specimens of material cut from the fractured blade were tested for

tensile strength, hardness, conductivity, and composition. All the tests indicated

values that were consistent with the specified composition for 7075-T73 aluminum

alloy.

1.16.2 Previous Failures of Similar Model Propellers

Prior to this accident, there were two failures of Hamilton Standard

composite-type propeller blades that were found to have resulted from cracks that

originated from inside the taper bore. The first event took place on March 13,

1994, on an Inter-Canadien

Aerospatiale-Aeritalia ATR 42 equipped with a

model 14SF propeller blade. The second event occurred on March 30, 1994, on a

Nordeste

Embraer EMB 120 equipped with a model 14RF blade. (Appendix C

contains details of the fractures).

The Transportation Safety Board of Canada (TSB) conducted an

investigation

of the Inter-Canadien event. TSP analysis indicated that forces

induced from the rotation of the three remaining blades resulted in propeller

imbalance and loads on the forward engine mounts that exceeded the ultimate

limits. This resulted in separation of the propeller and the RGB assembly from the

airplane. The RGB with the propeller hub, three complete blades, and the retained

portion of the fourth blade fell onto an ice-covered lake and was recovered during

the investigation. Indications were that the separated blade passed through the

fuselage and caused depressurization of the cabin. There were no injuries, and the

flightcrew accomplished a safe landing.

Inter-Canadien is a regional air carrier based in Montreal, Canada.

Nordeste Linhas Aereas Regionals S.A. is a regional air carrier based in Salvador,

Bahia, Brazil.

The TSB released the results of its investigation as Aviation Occurrence Report No.

A94Q0037 on February 28, 1995.

The Aircraft Accident Prevention and Investigation Center of Brazil

(CENIPA) investigated the Nordeste occurrence. CENIPA did not publish a

formal report; however, it provided documentation contained in a technical report

by Embraer that affirmed causal findings similar to the Inter-Canadien blade

separation. Embraer’s report indicated that during the Nordeste event, the

imbalance forces from the rotation of the three remaining blades resulted in

damage to the RGB. The remaining three blades and fourth blade stub were found

moved toward the feathered position (resulting in minimum aerodynamic drag);

and the propeller and RGB assembly remained within the nacelle area and were

partially attached to the airframe. There were no injuries, and this flightcrew also

accomplished a safe landing.

Laboratory examination of the failed blades indicated the presence of

chlorine-based corrosion pits in both instances. The chlorine source was traced to

a bleached cork installed in the taper bore to retain the lead balance wool. These

findings were corroborated by Hamilton Standard engineers and the FAA.

In addition to the Inter-Canadien and Nordeste propeller blade

fractures that were related to taper bore corrosion, on August 3, 1995, about 3

weeks prior to the ASA accident, there was an in-flight loss of a Hamilton

Standard Model 14RF-9 propeller blade that was not related to taper bore

corrosion. The propeller was installed on a Luxair

EMB-120 airplane that was in

the final approach to landing when the right propeller and portions of the RGB

separated. Some of the separating components struck the airplane; the flightcrew

accomplished a safe landing, and there were no injuries. The Belgian Civil

Aviation Administration (CAA) conducted an investigation on behalf of the

Ministry of Transport of Luxembourg.

The investigation determined that one of

the four propeller blades had failed from a fatigue crack about 9 inches from the

butt end of the blade. It was found that the crack began on the outer surface of the

blade shank in an area of mechanical damage induced by a localized interference

condition between the blade spar and the foam mold, which occurred during the

manufacturing process.

Luxair is a regional air carrier based in Luxembourg City, Luxembourg.

The Belgian CAA released the “Final Report of Aircraft Accident” on July 5, 1996.

1.16.3 Blade Inspection and Repair - Actions Taken by Hamilton

Standard and the FAA

(See Table 1 for a timeline of significant events related to the 14RF-9

propeller.) Following the March 1994 blade failures, Hamilton Standard began an

immediate program to inspect ultrasonically all model 14RF, 14SF, and 6/5500F

propeller blades for evidence of cracks. Blades with rejectable ultrasonic

indications were returned to Hamilton Standard. Early in the process of inspecting

the returned blades, Hamilton Standard discovered that some ultrasonic

indications were caused by visible mechanical damage. Although no cracks were

found, the mechanical damage was in excess of what engineers thought was

acceptable. Hamilton Standard reviewed the shop practices and concluded that the

mechanical damage was a result of tools and techniques used during the

installation and removal of balance wool lead. As a result, Hamilton Standard

developed repair procedures to blend locally visible mechanical damage and

eliminate ultrasonic indications that had no associated cracks. This repair was

described in Hamilton Standard repair procedure PS960 (and was approved by the

FAA on April 8, 1994.) PS960 specified the following steps:

1) Visually inspect the blade taper bore for evidence of mechanical

damage. No unblended mechanical damage is allowed.

2) Locally blend mechanical damage to 50 times the repair depth.

Repair limits are 0.010” maximum stock removal for the face area, 0.020”

maximum stock removal for all other areas, including end of taper bore.

When the blending is complete, no evidence of damage may remain.

Reference Figure 1 (page 3) for definition of face area at any taper bore

location.

3) Inspect repairs using a borescope with a 1:1 magnification to

verify blending to the above requirements. Surface finish of repair area

must be 63 RMS.

4) Perform an ultrasonic inspection of the blade taper bore area.

5) WARNING; CONVERSION COATING IS POISONOUS TO

EYES, SKIN, AND RESPIRATORY TRACT. USE SKIN AND EYE

PROTECTION. MAKE SURE THE TIME YOU USE IT IS THE

MINIMUM NECESSARY. MAKE SURE THE AREA HAS A GOOD

FLOW OF AIR.

6) Apply “PS960” to the face and camber side of each blade with

white stenciling ink in accordance with stenciling procedures provided in

the applicable Component Maintenance Manual.

With a brush, touch up all areas repaired per the above procedure

with a coating that agrees with MIL-C-5541, Class 1A. Allow to cure 24

hours.

NOTE: Alodine 600 is recommended because it is without cyanide, but

Alodine 1200 or 1201, or any material which agrees with MIL-C-5541,

Class 1A is satisfactory.

Soon thereafter, it was determined that the cork in the taper bore

contained chlorine residue that could cause corrosion in the taper bore. As a

result, PS960 was amended by PS960A to include procedures to eliminate the

taper bore cork and to replace it with a sealant. PS960A was approved by the

FAA on April 18, 1994.

Concurrent with the development of the PS960A repair procedure,

Hamilton Standard was also developing a series of alert service bulletins (ASB) to

address the problem of cracks originating from inside the taper bore in the model

14RF, 14SF, and 6/5500F blade spars. The bulletins called for a one-time, on-

wing ultrasonic shear wave inspection to be performed by level II

Hamilton

Standard employees or contractor inspection teams to detect abnormalities in the

blade taper bore. Blades rejected for ultrasonic indications above specified limits

found during the on-wing inspection were to be removed from service and sent to

Hamilton Standard Customer Support Centers. Upon receipt of the blades,

Hamilton Standard inspectors were certified according to the American Society for

Nondestructive Testing (ASNT) or the Hamilton Standard, FAA-approved equivalent.

According to ASNT, an NDT Level II individual is qualified to set up and calibrate equipment,

and to interpret and evaluate results with respect to applicable codes, standards, and

specifications. The NDT Level II is thoroughly familiar with the scope and limitations of certain

NDT methods, and guides and performs on-the-job training of trainees and NDT Level I

personnel. The NDT level III individual is familiar with other NDT methods and is capable of

training and examining NDT Level I and II personnel for certification in those methods.

TABLE 1

SIGNIFICANT EVENTS RELATED TO

HAMILTON STANDARD 14RF/SF SERIES PROPELLERS

Previous Accident(s): Blade Separation

Date Company Airplane Type

3/13/94 Inter-Canadien ATR-42

3/30/94 Nordeste EMB-120

Inspection and Repair Action

Date Document Reason Action

4/8/94

4/18/94

Hamilton Standard

PS960, as revised by

PS960A. (FAA

approved procedure.)

Mechanical damage

and chlorine deposits

found in taper bores.

-Visual inspection for

mechanical damage.

-Blend mechanical

damage.

-Remove cork and

replace with sealant.

4/18/94 Hamilton Standard

ASB 14RF-9-61-A66

Inter-Canadien &

Nordeste blade failures

due to fatigue cracking

-One-time on-wing

ultrasonic inspection to

detect abnormalities in

taper bore

- If rejectable

indications are found,

remove from service

and return to Hamilton

Standard

4/27/94 Hamilton Standard

Internal Memorandum

To document decision

to use 960A repair to

eradicate UT

indications caused by

peaks of shotpeen

impressions

-Was interpreted as

expansion of PS960A

blending repair to

include blades without

mechanical damage.

5/2/94 AD 94-09-06 -Required one-time UT

inspection for cracks in

taper bore (within 45

days) in accordance

with ASB 14RF-9-61-

A66.

- If cracks are found,

replace propeller.

ASA Accident Blade History

5/19/94 ASA accident blade inspected on-wing per AD and removed from service.

6/7/94 ASA accident blade inspected at Hamilton Standard, no visible faults found,

blend repaired per PS960A.

Inspection and Repair Action (Continued)

Date Document Reason Action

8/29/94 Hamilton Standard ASB

14RF-9-61-A69

(Revised10/5/94)

Continuing

airworthiness

-Repeat ultrasonic insp.

every 1,250 cycles for

specified unpeened

blades, or do improved

visual inspection per

Safety Board 14RF-9-61-

A70, or return to

Hamilton Standard.

- In all cases, return

rejected blades to

Hamilton Standard

-Remove cork (if still

installed).

- ASA blade exempted.

8/29/94 Hamilton Standard SB

14RF-9-61-A70

Continuing

airworthiness

-(For unpeened blades

only) Improved taper

bore visual inspection

with borescope photo

and mold transfer.

- If pits found, remove

from service or return to

repair facility.

9/1/94 Component

Maintenance Manual

(CMM)

No. 61-13-04

Detect corrosion, if

none then repair.

Eliminates PS960A

-Improved taper bore

cleaning & visual

inspection (cracks, pits,

and mechanical damage)

and FPI.

- Shotpeen.

3/ 23/95 AD 95-05-03 -Required SB 14RF-9-

61-A69 Rev 1, and SB

14RF-9-61-70 (by

12/31/97) and provided

terminating action to

repetitive inspection.

Accident(s): Blade Separation

Date Company Airplane Type

8/3/95 Luxair EMB-120

8/21/95 ASA EMB-120

Further Inspection and Repair Action

Date Document Reason Action

8/25/95 NTSB

recommendations

ASA Accident Recommended re-

inspection of reworked

blades; vibration and

loads survey; review of

requirements for

shotpeened taper bores.

8/25/95 TAD 95-18-51 ASA Accident and

NTSB

recommendations

Required re-inspection

of reworked blades.

9/30/95 14RF-9-61-86, Rev

4,11/9/95

LUXAIR blade shank

failure due to fatigue

cracking

Shank on-wing

ultrasonic inspection,

for “N” blades

11/9/95 14RF-9-61-A90

ASB

same as above Shank off-wing

ultrasonic inspection,

for “M” blades

11/16/95 AD 95-24-09 -Require SB 14RF-9-

61-A86 Rev 4 or SB

14RF-9-61-A90.

12/15/95 14RF-9-61-A91

ASB

Revised fracture

mechanics info. and risk

assessment of blade and

shank fatigue cracks.

-New off-wing

ultrasonic insp., lead

removed, each 500

cycles.

12/18/95 14RF-9-61-A95

ASB

Rev 1, 12/18/95

same as above -New on-wing

ultrasonic insp., with

balance lead.

1/19/96 AD 96-01-01 -Require ASB 14RF-9-

61-A91 or ASB 14RF-

9-61-A95.

Terminating Action

Date Document Reason Action

3/6/96 14RF-9-61-A94

ASB

terminate recurrent

ultrasonic inspections.

-Repair taper bore,

rework to like new.

4/24/96 AD 96-08-02 -Require ASB 14RF-9-61-

A94 taper bore

repair by 8/31/96.

Hamilton Standard performed a borescope inspection and initiated the repair

process, if warranted. The FAA mandated the inspection described in the ASBs

by AD 94-09-06, effective on May 2, 1994. The AD required that blades with

ultrasonic indications above 50 percent be removed from service. A rejectable

ultrasonic indication was found on the accident blade, and it was removed from

service on May 19, 1994. The accident blade was one of 490 rejected blades that

were sent to Hamilton Standard for further evaluation and possible repair.

Hamilton Standard customized inspection and repair instructions set

forth in a shop traveler form (Rock Hill Flow Traveler - Form Number RH243)

were used to define the taper bore inspection and repair actions for blades rejected

as a result of the field on-wing ultrasonic inspections. The shop traveler form

required that the taper bores of the returned blades be ultrasonically inspected

again to verify the unacceptable rejectable indication. The taper bores of the

blades were then to be cleaned and borescope inspected for evidence of cracks,

corrosion, pits, and other flaws. None of the returned blades were discovered to

be cracked. Corrosion was identified in approximately 13 percent of the blades;

these blades were set aside for further analysis by Hamilton Standard engineering.

Some of these blades were subsequently cut up or were otherwise destructively

tested; others were used to develop further testing and repair processes. None of

the blades with confirmed corrosion were returned to service.

After PS960A and the alert service bulletins were issued, Hamilton

Standard discovered that a small percentage of the returned blades with ultrasonic

indications did not have observable corrosion, mechanical damage, or cracks in

the taper bore. All such blades identified at that time had shotpeened taper bores.

Hamilton Standard determined that the roughness inherent in shotpeened surfaces

could in some cases also generate a rejectable ultrasonic indication. According to

Hamilton Standard engineering managers, as a measure to further reduce the

number of apparently unsubstantiated ultrasonic indications and to return these

blades to service, Hamilton Standard engineering personnel decided that the

procedures set forth in PS960A for blending areas of mechanical damage could

also be used to blend the area surrounding the ultrasonic indication inside the taper

bore of shotpeened blades, even though there was no associated mechanical

damage. This decision to extend the applicability of PS960A was discussed and

The shop traveler form that was in use at the time the accident blade was inspected and

approved by the engineering manager at Rock Hill on May 13, 1994 (approximately 1 month

before the accident blade was received there).

authorized in conference calls that included engineering managers of the three

Customer Service Centers and Hamilton Standard engineering. It was

implemented without the knowledge or approval of the FAA or the DER.

According to Hamilton Standard management, the authorization to

use PS960A in this manner was confirmed through an internal memorandum,

dated April 27, 1994, from the Manager of Operations Engineering to the three

Customer Service Center engineering managers, stating:

Subject: Blade U. T. [ultrasonic] Inspection

Per direction from [the head of Project Engineering] you should

handle blades returned from the field as a result of U.T. inspection

as follows:

1) Perform a U. T. inspection. Record results on the ASB form

except that this form should have the [applicable location]

written at the top of the form. Forward the form to HSD

Service via FAX.

2) Rework blade per PS960A. Perform a U.T. inspection and

record the results on the modified form as described in para. (1).

3) Ship acceptable blades. Hold rejected blades until further

notice.

In a letter to the Safety Board dated March 5, 1996, Hamilton

Standard indicated that the intent of this memo was to document the decision to

use PS960A “to eradicate false ultrasonic positives being caused only by

superficial irregularities,” specifically, to “remove tool marks or the peaks of shot

peen impressions.”

On August 29, 1994, Hamilton Standard issued an additional series of

ASBs and SBs for unshotpeened blades in the 14RF, 14SF, and 6/5500/F series

with procedures for repeating the on-wing ultrasonic inspection every 1,250 cycles

or, in the alternative, accomplishing a borescope inspection for pits. Rejected

blades were to be returned to Hamilton Standard for inspection (including FPI)

and repair according to the new maintenance procedure

in the Component

Maintenance Manual

(CMM), which superseded the procedures in PS960A. The

1,250 cycle interval was based on the minimum detectable flaw using the

ultrasonic inspection technique and the operating time to failure of the Nordeste

and Inter-Canadien blades.

Shortly after these ASBs were issued, the FAA issued AD 95-05-03,

effective on March 23, 1995, referencing the Hamilton Standard SBs and ASBs

and requiring that blades be ultrasonically inspected at an interval of 1,250 cycles

or, alternatively, that a borescope inspection of the taper bore be performed. The

AD provided appropriate ASB/SB references. If the borescope inspection found

no corrosion pits in the taper bore, the blade could be returned to service. AD 95-

05-03 also contained a provision that a return to service following the results of a

satisfactory borescope inspection constituted terminating action to the requirement

for recurrent ultrasonic inspections every 1,250 cycles.

Following this accident, the Safety Board made several urgent

recommendations resulting in additional ADs. (See Sections 1.18.1) Other

postaccident actions taken by Hamilton Standard and the FAA are discussed in

Section 1.18.2.

1.16.4 The Failed Propeller, Information, and Service History

The fractured propeller blade from N256AS was model 14RF-9, part

number RFC11M1-6A, serial number 861398, manufactured with an

unshotpeened taper bore in 1989 by the Hamilton Standard Division, United

Technologies Corporation, Windsor Locks, Connecticut. The 14RF-9 model blade

is certificated only for the EMB-120 airplane.

Repair 4-25 in the CMM required the removal of the bore plug, cork (if installed), and

lead wool, followed by a cleaning, white light borescope and FPI inspections. For unpeened

blades: If a blade had damage less than 0.005 inch, with no previous blending per PS960A, then

the blade was shotpeened, balanced and marked +A. If the blade had damage less than 0.005

inch, but it was previously blended per PS960A, then the blade was shotpeened, balanced and

marked +B. If the damage was greater than 0.005 inch, but less than 0.020 inch (0.010 inch on

the blade face), then the blade was reamed, shotpeened, balanced and also marked +B. If the

damage was greater than 0.020 inch (0.010 on the blade face), then the blade was reamed,

balanced and marked +C. The process was similar for shotpeened blades.

The CMM is the FAA-approved maintenance manual that contains instructions for

continued airworthiness of Hamilton Standard propellers.

The fractured blade had accumulated a total of 14,728 operating

hours and 5,182 hours since overhaul. The overhaul was accomplished by

Hamilton Standard customer service technicians at East Windsor on April 7, 1993.

At that time, the blade had accumulated 9,546 hours.

Records indicated that only

routine maintenance actions were necessary at the time of overhaul.

On April 20, 1993, upon return to ASA from overhaul, the blade was

installed on an airplane (not the accident airplane) where it remained until May 19,

1994. On that day, the blade received an on-wing ultrasonic inspection of the

taper bore by a Hamilton Standard contract inspector in accordance with AD 94-

09-06. The blade was rejected for a 60-percent, full-scale height

ultrasonic

indication.

The blade was removed from service and returned to Hamilton

Standard facilities at East Windsor, Connecticut. It was subsequently shipped to

Hamilton Standard’s Customer Support Center at Rock Hill, South Carolina, for

inspection and repair.

According to the shop traveler form for the accident blade, an

ultrasonic inspection of the accident blade on June 7, 1994, confirmed the

rejectable indication, with a reading of 52 percent full-scale height. Following this

ultrasonic inspection, the lead wool was mechanically removed from the taper

bore and the hole was examined with a white-light borescope for evidence of

corrosion, pits, or cracks. In the space provided on the shop traveler form for the

“results” of this inspection, the technician recorded, “No visible fa[u]lts found,

blend rejected area.”

The shop traveler form reflects that he then blend-repaired

the taper

bore “damage” with aluminum oxide sanding tools, using the procedures of

PS960A. The technician, who was not an FAA-certificated mechanic, stated that

The inspection limit for 14RF-9 propeller blades is 9,500 hours. ASA had FAA

approval to fly in excess of the required inspection time by as many as 500 hours.

Pursuant to criteria set forth in Hamilton Standard ASB 14RF-9-A66 (incorporated by

reference in AD 94-09-06), indications of 50 percent full-scale height and above, as viewed on a

cathode ray tube screen, were rejectable. Indications of 40 to 50 percent were “reportable” on the

inspection record but were not cause to reject the blade.

Blending in this context means grinding or sanding the surface to remove a small

amount of material containing an imperfection and then restoring the surface finish to a condition

equal to the surrounding area.

he was permitted to perform and sign off the work that he was qualified to

perform. The technician, as an employee of Hamilton Standard’s Rock Hill blade

repair facility, which is an FAA-certificated repair station under 14 CFR Part 145,

is not required to be a certificated mechanic to work on the propeller blades. In

the shop traveler form for the accident blade, the instructions to blend the taper

bore also specify that the surface finish should be 63 RMS. The block on the form

was initialed and dated by the technician, but an adjacent block that would have

signified a second inspection by a repairman

was blank. Although the shop

traveler form for the ultrasonic inspection after the blending on the accident blade

showed that there were no reportable or rejectable indications, the shop traveler

form did not show that the blade had received a final inspection after the work was

completed. However, Hamilton Standard provided ASA with an FAA Form 8130-

3, Airworthiness Approval Tag, indicating that the PS960A repair had been

completed.

Subsequent examination by the Safety Board indicated that about

0.002 inch of material was removed from the taper bore. The shop traveler form

indicates that the blade passed a postrepair ultrasonic inspection.

Because no

defects were found during the June 7, 1994, borescope inspection, and the blade

was marked with “PS960A” (indicating accomplishment of the repair), the blade

was exempt from the requirements of AD 95-05-03 (effective on March 23, 1995)

for recurrent taper bore ultrasonic inspections and enhanced borescope inspection.

Investigators attempted to establish how the decision was made to

blend the area of the unacceptable indication in the taper bore of the accident

blade, even though there was no visible mechanical damage and PS960A did not

explicitly require or authorize blending of a blade taper bore that was free of

visible mechanical damage. During an interview with the technician who

performed the taper bore repair, he stated that he had been told that the

shotpeening of the taper bore could cause “false” ultrasonic indications. Further,

he stated that he understood, based on his training in the repair by the Rock Hill

A repairman, as defined in 14 CFR, Parts 65.101 and 65.103, is recommended for

certification by the repair station and is then certificated by the FAA. He must have at least 18

months of experience on the specific task, have completed specialized training, and may

supervise maintenance of aircraft components by the repair station.

The shop traveler form indicated that a blade that “failed” the postrepair, ultrasonic

inspection could be reblended. If the blade “failed” the ultrasonic inspection after the second

attempt to blend the rejected area, it was to be sent to Hamilton Standard’s Windsor Locks

facility.

facility Engineering Manager, that it was acceptable to use the PS960A process to

blend out unexplained ultrasonic indications for blades with unshotpeened taper

bores, as well as those with shotpeened taper bores. He said that he recognized the

difference in surface finish between shotpeened and unshotpeened blades. He also

stated that if he came across something that he did not understand or recognize, he

would not hesitate to seek assistance from the facility Engineering Manager. The

technician further stated that he had blended about 10 propeller blades without

shotpeened taper bores that had ultrasonic indications but no visible damage.

During an interview on October 19, 1995, the Rock Hill Engineering Manager

stated that the April 27, 1994, memorandum covered both shotpeened and

unshotpeened blades. However (as already noted in Section 1.16.3), in a letter to

the Safety Board dated March 5, 1996, Hamilton Standard indicated that the

memorandum was intended to document the authorization to use the PS960A

blend repair to “remove tool marks or the peaks of shot peen impressions.” The

letter further stated that “there was no discussion of how to handle blades that had

not been shot peened.”

After the PS960A blending repair, the accident blade was rebalanced

and the taper bore was sealed with protective material. It was also determined that

some additional repair was necessary to the composite surface features of the

blade. The blade was returned to Hamilton Standard’s East Windsor, Connecticut,

facility to complete that work. The blade was shipped back to ASA on August 30,

1994, and was reinstalled on the left propeller assembly of the accident airplane on

September 30, 1994. It remained there until the accident. At the time of the

accident, the blade had accumulated 2,398.5 hours and 2,425 cycles since the

Hamilton Standard repair at Rock Hill.

1.16.5 Results of 14RF-9/EMB-120 Stress Survey

To reduce cabin noise during ground operation, the EMB-120 aircraft

was certificated with model 14RF-9 propeller assembly designed to rotate at a

relatively low ground idle revolutions per minute (rpm), between 50 and 65

percent. To operate successfully in the 50 to 65 percent rpm range without

overstressing the propeller, the design had to avoid the coincidence of any blade

resonant frequencies

with any excitation frequencies.

To accomplish this, the

The resonant frequency of any vibration is the naturally occurring frequency at which

the blade will vibrate when excited. To avoid excessive vibration and overstressing of the

upper rpm limit of 65 percent for ground operation was set below the resonant

frequency of the first flat-wise mode of vibration

of the 14RF-9 blade of

approximately 70 percent rpm. Although all 14RF blades have a resonant

frequency at approximately 70 percent rpm, because of small differences in blade

construction, there is some variation in the exact resonant frequency from blade to

blade.

The exact resonant frequency of the accident blade could not be

determined because it had separated and was damaged.

The Hamilton Standard 14RF-9 blade for the EMB-120 was designed

so that its first flat-wise resonant frequency would not coincide with the 2P

excitation frequency during ground operation. However, the coincidence or close

proximity of the ground rotational speed and the first flat-wise resonant frequency

can place the propeller in a resonant condition and results in undesirable vibratory

stress. (See Figure 4 for an illustration of the vibratory modes of the 14RF-9

propeller.)

propeller, propeller design practice requires that the propeller spend only a minimal amount of

time in an rpm range that corresponds to a resonant frequency.

Excitation frequencies are created by aerodynamic loads that act on a propeller. Some

aerodynamic loads on a propeller are cyclical. The frequency of these cyclical loads are

multiples of propeller rpm. The frequency of the first cyclical aerodynamic load (1P) is equal to

propeller rpm; that is, one cycle per revolution.

The first flat-wise mode of vibration is the lowest frequency vibration. (There are

multiple orders of vibration, and each has an associated frequency called a resonant or natural

frequency.)

A resonant frequency of a propeller blade is a function of the blade’s rigidity, mass

distribution and, to some extent, retention stiffness.

The frequency of the second aerodynamic load (2P) acts on all propeller blades twice

per revolution because the blade senses the wing leading edge as it rotates past it. The 2P

excitation frequency is always present but is most pronounced during ground operation in a

tailwind or quartering tailwind condition.

TECHNOLOGIES

wLm

‘W-

CAMPBELL DIAGRAM FOR

EMB120

JAV-9-1 2-95

HAMILTON STANDARD

14RF-9

PROPELLER

70-

Ill

APPROXIMATE

RESONANT RPM

BAND FOR Ist MODE

("REACTIONLESS")

‘~

,Allowable’

Range for

Static

Pi”OPELL:#

PERC:~T

SPEE:O(l 00%

1300

RPh~

90 1

100

200

300 400 500 600

700

800

900

1000

1100

1200

1300

PROPELLER RPM

Figure 4.--Vibratory modes of the

14RF-9

propeller.

When the first flat-wise vibration frequency and the 2P excitation

frequency coincide, a four-bladed propeller, such as the 14RF-9, vibrates in the

reactionless

mode. The reactionless mode of vibration is characterized as when

two blades of a four-bladed, rotating propeller, 180 degrees apart, reach a negative

stress peak at the same time that the other two blades reach their positive stress

peak. This type of propeller vibration is called reactionless because the bending

loads of the four blades are canceled at the propeller hub which consequently

transmits little or no vibratory loading to the propeller mounting structure.

As originally certificated, the EMB-120 Airplane Flight Manual,

Maintenance Manual, and a cockpit placard stated that propeller rpm (Np)

above

65 percent should be avoided during ground operations with the aircraft stationary,

and with rear quartering winds in excess of 10 knots, except for short duration

transitions.

The 1983 and 1985 certification testing of the EMB 120 with the

model 14RF propeller was accomplished using an instrumented, new propeller

blade. The original stress surveys, which included flight and ground

tests, and an

additional ground test following the ASA accident, revealed that the resonant

frequency of the first flat-wise mode varied from approximately 70 percent to 76

percent. Because the actual resonant frequency of a propeller must be determined

during ground testing, and because propeller assemblies are not measured

The Safety Board investigated another accident on April 19, 1993, in Zwingle, Iowa, in

which a Mitsubishi MU-2B-60 airplane experienced an in-flight loss of a propeller blade and

collided with terrain (NTSB/AAR-93-08). All of the occupants were killed, and the airplane was

destroyed. The Safety Board attributed the loss of the propeller to a reduction in the fatigue

strength of the hub material, combined with exposure to higher-than-normal cyclic loads during

ground operations when the propeller vibrated in the reactionless mode. The investigation

revealed that during certification testing of the Hartzell HC-B4 propeller on the MU-2B airplane,

a reactionless mode of vibration was identified with the peak stress occurring at a propeller speed

of 1,079 rpm. As a result, the propeller was prohibited from continuous operation on the ground

below 1,145 rpm. The Safety Board did not find evidence that the test was repeated using

propeller blades altered to conform to the minimum dimensions specified in the repair limit

criteria of Hartzell’s HC-B4 propeller maintenance manual.

Propeller rpm expressed as a percentage of the maximum operational limitation.

The test airplane was subjected to the propeller wash of a second airplane to simulate

normal ground operations in a quartering tailwind. This was accomplished to determine the high

stress peaks of the propeller blade associated with the 2P excitation frequency within the normal

propeller rpm range.

separately, the actual resonant frequency of the accident blade could have been

slightly different than the frequencies measured in these tests.

Following the accident, Hamilton Standard conducted an independent

blade stress survey in the latter part of 1995 and presented its report as part of a

technical presentation to Safety Board investigators on February 2, 1996 (see

Appendix D for Summary of Results, Conclusions and Recommendations). The

1995 stress survey also consisted of a flight and ground test, using four

instrumented blades on both engines with two taper bore configurations. Of the

four blades, one blade had zero service time, and the other three blades had

approximately 12,000 service hours. The survey concluded that only small

differences

from previous stress level surveys were observed, including those

from the original certification and the postaccident test referenced above. The

highest vibratory stress measured in flight was reported to be about 6,000 pounds

per square inch (psi). The highest ground running peak stress was reported to be

about 18,000 psi. This occurred with quartering tailwinds when the propeller

passed through the range corresponding to vibration of the blade twice per

revolution in the first flat-wise mode (2P/1f).

After the ground tests were completed, Hamilton Standard initiated

actions to modify the EMB-120 Airplane Flight Manual and Maintenance Manual

Limitations Sections with revised language to reduce the maximum propeller Np

limit during ground operation from 65 to 60 percent. On December 20, 1995,

airworthiness authorities in Brazil and the FAA approved temporary revisions to

the language as follows:

Airplane Flight Manual: Condition Levers must be in MIN RPM

position during all ground operations, except when cleared for

takeoff or during landing roll. Power Levers must remain at or

below Flight Idle during all ground operations, except for brief

(approximately 5 seconds) excursions as needed to maneuver the

airplane.

The vibration and loads survey was conducted by Hamilton Standard on the 14RF

propeller on an Embraer EMB-120 airplane. A possible small downward shift was discovered in

the 2P resonant frequency relative to the 1985 certification test data. The downward shift was

attributed by Hamilton Standard to normal wear and mass properties in the propeller blade from

normal operation.

CAUTION: GROUND OPERATION ABOVE FLIGHT IDLE

SIGNIFICANTLY INCREASES PROPELLER STRESS UNDER

CERTAIN ADVERSE WIND CONDITIONS (E.G., TAILWINDS

OR REAR CROSSWINDS). OPERATION IN THIS RPM

RANGE MUST BE AVOIDED TO THE MAXIMUM EXTENT

PRACTICABLE.

Maintenance Manual: Propeller System Operating Limitations, To

prevent excessive propeller stress, do not operate above 60% Np

unless the wind is less than 18.5 Km/h (10 Knots) or the airplane

is headed into the wind + or - 45 degrees. Wind direction must be

monitored locally at the run-up site.

1.17 Organizational and Management Information

1.17.1 Hamilton Standard Division, United Technologies Corporation

Hamilton Standard Division, Windsor Locks, Connecticut, has

produced aluminum blade propellers with taper bores since 1958. More than

10,000 such blades are in use on C-130 and P-3 military airplanes. Composite

blade propellers were introduced for military airplanes in 1974. Their use was

expanded to regional (commuter) airplanes in 1978. Since that time, regional

airplane propellers equipped with more than 15,000 blades that are fabricated with

an aluminum spar and an aerodynamic shell of composite materials have

accumulated in excess of 17 million flight hours.

A customer support facility for regional propeller inspection and

repair was opened in East Windsor, Connecticut, in 1989. Hamilton Standard’s

engineering support continued to reside with the manufacturing facilities at

Windsor Locks.

From 1989 through 1990, regional propeller inspection and repair

expanded to facilities in Long Beach, California, and Maastricht, the Netherlands.

In 1993, the regional propeller customer support center was moved from East

Windsor, Connecticut, to a facility in Rock Hill, South Carolina.

1.17.2 Hamilton Standard Propeller Customer Service Center

Selected engineering and management personnel from the Hamilton

Standard Division, Windsor Locks, were transferred in 1993 to the new regional

propeller Customer Support Center in Rock Hill. Management at the Rock Hill

facility hired technicians from a large pool of local area applicants and trained

them to perform various propeller blade repairs. The facility was certificated by

the FAA to begin repairing propeller blades in February 1994.

As a result of the two propeller blade failures in March 1994,

and

the resulting ultrasonic inspections of the taper bore mandated by the FAA

May 1994, there was a sudden increase in the number of propeller blades requiring

inspection and repair in May and June 1994. The accident propeller blade was one

of the blades returned to Hamilton Standard as a result of the inspection. The

taper bore inspection and repair of the accident blade was performed at Hamilton

Standard’s Customer Support Center in Rock Hill between June 7, and June 9,

1994.

According to Hamilton Standard work records from May and June

1994, the technician who performed the taper bore inspection and repair of the

accident blade worked between 8 and 26 hours of overtime each week, in addition

to his normal work week.

1.17.2.1 Employee Training at Hamilton Standard

Customer Service Center

According to the Engineering Manager in Rock Hill, new technicians

at the facility received approximately 250 hours of training on a specific propeller

blade repair before they are permitted to make repairs without direct supervision.

However, the technician who made the taper bore repair to the accident blade was

transferred from the nickel sheath and fiberglass repair area of the shop and had

received 250 hours of training in connection with that position. In addition, he

received approximately 90 hours of additional on-the-job training (OJT) on the

taper bore repair. This technician had a background in automotive repair. His

additional OJT training on the taper bore repair procedure was administered by the

Details of the two previous failures are contained in Section 1.16.2.

AD 94-09-06 required an ultrasonic inspection of the taper bore. (See Section 1.16.3

for details on the AD.)

facility Engineering Manager, who had other responsibilities in addition to

employee training.

The investigation revealed that during the time the accident blade was

being repaired, the technicians at the service centers had not been provided a

photograph or model illustrating the appearance of corrosion or cracking in a taper

bore. Neither the Engineering Manager nor the technician who performed the

inspection and repair of the accident blade had ever seen a crack in the taper bore

of a blade. Several months after the accident blade was inspected and repaired, all

the service center technicians were provided with an enlarged color photograph of

taper bore corrosion, a model and an instructional video.

1.17.3 Designated Engineering Representative

The FAA annually appointed several Hamilton Standard engineering

employees based in Windsor Locks to serve as Designated Engineering

Representatives (DER) in propeller systems to approve certain engineering

information.

Those DER duties normally represented about 20 to 30 percent of

his or her workload. There were nine DERs certificated to support propeller

systems at the time of the accident. Under an agreement with the FAA, Engine

and Propeller Directorate, DERs have authority to approve certain engineering

changes, repairs, service bulletins and maintenance manuals. However, the

requirement for direct FAA approval was retained for changes affecting critical

parts or single point failure components.

After the two blade failures in March 1994,

a Hamilton Standard

DER

was in contact with the FAA, Engine and Propeller Directorate, several

times a week during the development of the ultrasonic inspection requirements

that were set forth in the ASB subsequently mandated by AD 94-09-06, and the

taper bore repair procedures contained in PS960 and PS960A. The DER was

In accordance with 14 CFR 183.29(f), a propeller engineering representative may

approve engineering information relating to propeller design, operation, and maintenance, within

limits prescribed by the Administrator of the FAA, whenever the representative determines that

information complies with applicable FARs.

See Section 1.16.2 for details.

The DER who was the central figure in the activity related to this accident was initially

appointed in October 1992. He had 16 years of experience at the Windsor Locks propeller

facility. He holds a Bachelor of Science degree in astronomy and a Master’s Degree in business

administration. He also attended a university-level course in aircraft accident investigation.

involved in developing the ASB and PS960/960A. The DER was familiar with the

issue of taper bore corrosion, the formation and functions of the Rock Hill

Customer Service Center, and the white light borescope inspection technique used

to search for evidence of cracks and corrosion, and to reject blades for engineering

evaluation. After the accident, he told investigators that based on his knowledge

of the March 1994 accidents, he believed the white light borescope inspection

would be adequate to enable technicians to detect corrosion pits in the taper bore

of the size that led to the cracks that caused the two previous failures.

The DER stated that he was not aware of the discussions between

engineering managers at Windsor Locks and the Customer Service Centers in

which it was decided to extend the taper bore repair procedures outlined in

PS960A beyond their intended use (to blend visible mechanical damage in taper

bores) to blend ultrasonic indications that were not related to visible mechanical

damage. He said that this lack of coordination from the engineering department

was atypical.

1.17.4 FAA Certification Engineer

The FAA certification engineer

responsible for Hamilton Standard

regional propellers reviewed the fracture analysis of the two blade fractures in

March 1994. He determined that the FAA should retain direct approval authority

for the data and the inspection and repair techniques used for PS960/PS960A

because, based on the two blade failures that resulted from taper bore corrosion, he

considered the taper bore to be a critical part.

The Certification Engineer also

required that the repair for mechanical damage in the taper bore be accomplished

only at Hamilton Standard repair facilities.

The FAA certification engineer had been with the FAA Engine and Propeller

Directorate for 9 years at the time of the accident. He holds a Bachelor of Science degree in

mechanical engineering. He has 30 years of government experience, primarily with U.S. Navy

programs. His experience includes propellers for surface ships and submarines.

FAA Order 8110.37A (DER Guidance Handbook), Paragraph 14.h.(1)states that “The

DER must obtain specific authorization from the appointing ACO [Aircraft Certification Office]

prior to initiating approvals for repairs or alterations. An authorized DER may approve technical

data for major repairs and alterations without first notifying the project ACO, except when the

part is critical or life limited. For critical or life limited parts, the DER must contact the project

ACO for guidance.”

The Certification Engineer (like the DER) stated that he was not

aware that blending procedures outlined in PS960A (originally approved as a

repair for areas with visible mechanical damage) had been extended by Hamilton

Standard engineers as an authorized repair to areas in which ultrasonic indications

were present without visual evidence of mechanical damage.

1.18 Additional Information

1.18.1 Safety Board Recommendations

The Safety Board issued the following three Safety Recommendations

to the FAA on August 25, 1995, which was 4 days after the accident:

A-95-81 (Class I, Urgent Action)

Immediately implement the ultrasonic inspection program on

Hamilton Standard propeller blades cited in paragraph (a)(2) of

airworthiness directive (AD) 95-05-03, irrespective of prior

compliance with paragraph (d) of the AD. Require the initial

inspection before further flight on any propeller blades that have

accumulated 1,250 cycles since the last ultrasonic inspection or

since the visual and borescope inspection required by paragraph

(d) of the AD.

A-95-82 (Class II, Priority Action)

Conduct a vibration and loads survey and analysis of the propeller

installation on the Embraer EMB-120 airplanes with applicable

Hamilton Standard propellers throughout the ground and flight

operating range of the engine with specific consideration for the

effects [that] propeller in-service wear, maintenance, or other

changes may have on the resonant frequencies. Based on the

findings, broaden the survey and analysis to other installations as

appropriate.

A-95-83 (Class II, Priority Action)

Review the current overhaul and inspection requirements for all

Hamilton Standard 14 series propeller blades for which the taper

bore hole has been shotpeened to determine whether additional

inspections or maintenance should be required.

Regarding Safety Recommendation A-95-81, on August 25, 1995, the

FAA issued a telegraphic AD (T95-18-51) requiring that certain blades installed

on EMB-120 aircraft that had ultrasonic crack indications discovered as a result of

inspections required by AD 94-09-06 or AD 95-05-03, been reworked, and had

been returned to service, be removed from service within the next 10 flight cycles

and be replaced with serviceable parts. The telegraphic AD also required that

propeller blades installed on aircraft other than the EMB-120 meeting the same

criteria (ultrasonic crack indications reworked and blade returned to service) be

inspected ultrasonically for cracks within the next 10 flight hours and every 1,250

cycles thereafter. Any blades removed from service as a result of the AD could

not be returned to service. On January 16, 1996, the Safety Board classified

Safety Recommendation A-95-81 “Closed--Acceptable Alternate Action.”

Regarding A-95-82, on July 24, 1996, the FAA informed the Safety

Board that it conducted the requested vibration and loads survey. The FAA

reported that the survey substantiated the results of past stress surveys indicating

that no new high stress conditions were uncovered, and that it did not show that

further evaluations were needed on other installations. However, to limit propeller

exposure to known high vibratory stresses during ground operation, on December

27, 1995, the FAA issued AD 95-25-11, which requires installation of a placard

reading “Avoid Np Above 60% During Ground Operations,” and also requires

revising the EMB-120 Airplane Flight Manual and maintenance program to limit

the rotational speed of the propeller during ground operation. In its Final Rule,

published in the Federal Register, the FAA stated that the AD is considered “an

interim action until final action is identified, at which time the FAA may consider

further rulemaking.” Based on the FAA’s interim actions and the statement

contained in its announcement of the Final Rule, on November 15, 1996, the

Safety Board classified Safety Recommendation A-95-82 “Open--Acceptable

Response.”

Regarding A-95-83, on July 24, 1996, the FAA informed the Safety

Board that its review of the overhaul and inspection requirements for all Hamilton

Standard 14RF, 14SF and 6/5500/F blade designs for which the taper bore hole

has been shotpeened showed that additional action should be taken. On April 24,

1996, it issued AD 96-08-02, requiring repetitive ultrasonic inspection of the

blades until they are repaired and restored to their certificated strength. The repair

of EMB-120 blades was required to be accomplished by August 1996, and all

other blades will be repaired by February 1997. Since the actions of the FAA are

responsive to the intent of the recommendation, on November 15, 1996, the Safety

Board classified Safety Recommendation A-95-83 “Closed--Acceptable Action.”

On June 27, 1996, the Safety Board issued the following additional

safety recommendations to the FAA:

A-96-33

Conduct a design review of the Embraer EMB-120 flight data

recorder system, with emphasis on potentiometer failures, and

mandate design, installation, and/or maintenance changes, as

necessary, to ensure that reliable flight control data are available

for accident/incident investigation.

A-96-34

Require Embraer EMB-120 operators to perform a flight data

recorder (FDR) readout or a potentiometer calibration test per

section 31-31-00 of the EMB-120 Maintenance Manual every 6

months until FDR sensor design, installation, and/or maintenance

improvements are incorporated.

On September 5, 1996, the FAA responded that it had initiated a

design review focusing on the potentiometers and associated attaching hardware,

and would determine a course of action when the review is complete. The FAA

also stated that it would contact the manufacturer and coordinate the necessary

maintenance instructions with an appropriate inspection interval. Therefore, on

October 15, 1996, the Safety Board classified both A-96-33 and A-96-34 “Open--

Acceptable Response,” pending final action.

1.18.2 Postaccident Hamilton Standard and FAA Actions

As a result of the accident and Safety Board recommendations, on

August 25 and 28, 1995, the FAA issued AD T95-18-51 and AD 95-18-06,

respectively, which required that all blades installed on EMB-120 aircraft that, like

the accident blade, had been removed from service in accordance with AD 94-09-

06 (or 95-05-03) and been reworked and returned to service be immediately

removed from service, and it required an ultrasonic reinspection of all other 14RF,

14SF, and 6/5500/F blades (a total of approximately 15,000 blades) on a 1,250

cycle interval.

As a result of the Luxair accident, Hamilton Standard developed an

additional series of service bulletins that were mandated by the FAA in

Airworthiness Directive (AD) 95-24-09, issued on November 16, 1995. This AD

required an ultrasonic shear wave inspection on the propeller shank for cracks or

indentations. Four blades were removed from service as a result of this AD.

As the inspections mandated by AD 95-18-06 began to take place,

Hamilton Standard recognized a need for a detailed reevaluation of the

adequacy/appropriateness of this inspection procedure because the ultrasonic

inspections continued to reject blades with no apparent discrepancies on the taper

bore. This reevaluation showed that the minimum detectable size of a taper bore

crack varied widely. That is, it could be much greater (and, in some cases,

smaller) than previously believed. Hamilton Standard found that there were

differences in the calibration blocks

distributed for the inspection, and that

differences in the thickness of the surface composite layers and in the

transmitability of the ultrasonic beam through the aluminum spar could

substantially affect the inspection results.

Coincidentally, Hamilton Standard performed a risk analysis of the

failed blades to resolve inconsistencies in the crack growth data from the Inter-

Canadien, Nordeste, and ASA blade separations. Results from a NASA-developed

fracture mechanics FASTRAN program were integrated into the risk analysis.

Hamilton Standard concluded that for some blades of each model, the 1,250 cycle

inspection interval using the original ultrasonic technique provided insufficient

safety margins for the detection of taper bore cracks. Thereafter, Hamilton

Standard issued another series of ASBs for the 14RF, 14SF, and 6/5500/F blades,

containing a newly-developed ultrasonic inspection technique with a customized

calibration procedure for each blade and a reduced inspection interval. For the

14RF blades used on the EMB-120 airplanes, the inspection interval was reduced

to 500 cycles. The FAA issued AD 96-01-01, effective on January 19, 1996,

requiring the improved ultrasonic inspection technique and mandating the reduced

cyclic inspection interval.

In March 1996, Hamilton Standard issued ASB 14RF-9-61-A94 to

repair and restore the taper bore surface to its original surface finish (with

shotpeening) on all model 14RF-9 propeller blades installed on EMB 120

Test specimens manufactured with defects of known sizes used to adjust the threshold

of pass/fail for ultrasonic inspection equipment.

airplanes and on similar model blades installed on other commuter-type airplanes.

On April 24, 1996, the FAA issued AD 96-08-02, which required that all blades of

the affected propellers be repaired by specific end dates. The end date for the

14RF-9 model blade was August 31, 1996, and for other affected models is

February 28, 1997. The repair procedure included removal of a layer of spar

material from the taper bore, followed by an eddy current inspection, fluorescent

penetrant inspection (FPI), removal of an additional layer of material, a wall

thickness check, shotpeening, and the application of a corrosion-protective

coating. Hamilton Standard also conducted fatigue tests on a set of blades with

the minimum allowable remaining wall thickness between the taper bore and the

outside of the spar, thereby verifying the fatigue characteristics of blades repaired

in this manner. Once a blade is repaired, the requirements for ultrasonic

inspection are eliminated, and further visual inspections of the taper bore will take

place only as part of the major inspection that is required every 9,500 flight hours.

There are no calendar-based inspection requirements.

Hamilton Standard informed the Safety Board that all 14RF-9 blade

repairs were accomplished within the time limit established in the AD. The Safety

Board also notes that the Component Maintenance Manual (CMM) for all

Hamilton Standard model 14FR, 14SF, and 6/5500F airplane propellers (which

includes all propellers used in commuter operations), has been amended to include

a recurring inspection of the taper bore.

2. ANALYSIS

2.1 General

The flightcrew was trained, certificated and qualified to conduct the

flight, and the flight was conducted in accordance with applicable FARs and

company requirements. The flight attendant also was appropriately trained and

qualified. The flightcrew was in good health and held the proper FAA medical

certificates. There was no evidence that the performance of any crew member was

impaired by alcohol, drugs, or fatigue.

The airplane was maintained in accordance with applicable FARs and

company Operations Specifications. A review of the airplane’s maintenance

records and operating history did not reveal any maintenance discrepancies or

mechanical anomaly that would have either caused or contributed to the accident.

Evidence from the CVR, FDR, and examination of the powerplants,

reduction gear boxes, and propellers indicated that the engines were operating

normally during the flight until the loss of a major portion of one propeller blade

on the left engine. After the propeller blade separation, the combination of the

resulting loss of left engine thrust, increased drag from a deformed engine nacelle

and the three blades retained in the propeller hub, and added frontal drag from

external sheet metal damage, degraded airplane performance, preventing the

flightcrew from arresting the airplane’s descent or making rapid changes in its

direction of flight making a forced landing necessary. The Safety Board concludes

that because of the severely degraded aircraft performance, the flightcrew’s

actions were reasonable and appropriate during their attempts to control and

maneuver the airplane throughout the accident sequence and were not a factor in

this accident.

2.2 Analysis of the Propeller Blade Failure

The Safety Board concludes that one of the four blades from the left

engine propeller separated in flight because a fatigue crack that originated from

multiple corrosion pits in the taper bore surface of the blade spar propagated

toward the outside of the blade, around both sides of the taper bore, then reached

critical size. (See Section 1.16.1.)

Results of investigations conducted in two previous propeller blade

failures in 1994, one in Brazil with this model blade and the other in Canada with

a similar model blade, indicated that corrosion was produced when entrapped

moisture reacted with residual chlorine in a bleached cork used to retain the lead

wool in the taper bore hole of the propeller. The accident blade exhibited a nearly

continuous layer of oxide deposits on the initial 0.049 inch of the crack depth.

These deposits contained a substantial amount of chlorine. The Safety Board

found that the ASA propeller blade contained corrosion damage (pitting) in the

taper bore and the oxide layer in the origin area of the fatigue crack in the

separated ASA propeller blade, as did the two previous failed propellers. Because

the oxidizing condition and the cork containing chlorine were eliminated from the

accident blade per PS960A during the June 1994 repair at the Hamilton Standard

Customer Support Center in Rock Hill, oxide deposits would not be expected to

have formed during further crack propagation after the PS960A repairs.

Therefore, the extent of the oxide layer on the fracture (to a depth of 0.049 inch

below the surface) was indicative of the size of the crack at the time of the Rock

Hill repair activities.

2.2.1 The Accident Blade’s June 1994 Inspection, Repair, and Return

to Service

2.2.1.1 Inappropriate Use of PS960A Blending Repair

As discussed in Section 1.16.4, the technician who inspected and

repaired the accident blade first confirmed the rejectable ultrasonic indication, and

then visually examined the taper bore for evidence of corrosion, pits or cracks

using a white light borescope. He wrote on the shop traveler, “No visible fa[u]lts

found, blend rejected area,” and used the blending repair procedure set forth in

PS960A to remove the ultrasonic indication. The blended area was later found to

be the site of a crack originating in corrosion pits.

Apparently, Hamilton Standard engineering originally intended

PS960A only to remove possible sources of stress concentration by blend-

repairing mechanical damage (visible tool marks) within the taper bore of any

blade, without regard for whether the surface was shotpeened or not shotpeened.

The instructions in PS960A with regard to the surface finish of the taper bore

specifically stated, “No unblended mechanical damage is allowed.” The FAA

reviewed and approved the repair for this purpose. However, the use of PS960A

blending repair was expanded by Hamilton Standard engineering to blend the area

of ultrasonic indications even when there was no apparent mechanical reason

(visible tool mark) associated with the ultrasonic indication.

The Safety Board considered whether it was appropriate, from an

engineering perspective, for Hamilton Standard to extend the applicability of

PS960A beyond its original purpose (blending of mechanical damage), and to

authorize its use for removing ultrasonic indications caused by shotpeen

impressions. Surface irregularities created by shotpeening are, in effect, a form of

mechanical surface alteration, and the concept of blending mechanical damage is

not per se objectionable, so long as there are no cracks or other defects in the area

being blended.

Based on the prior blade separations (both of which involved

cracks originating from corrosion), Hamilton Standard had no reason to believe

that mechanical damage in taper bores was causing cracks. Therefore, the Safety

Board concludes that Hamilton Standard’s engineering decision to use the

PS960A blending repair to remove ultrasonic indications caused by a shotpeened

taper bore surface was technically reasonable.

Although the decision by Hamilton Standard engineers to extend the

applicability of PS960A to impressions in shotpeened taper bores was technically

reasonable, the procedure by which that decision was communicated to others

within Hamilton Standard was deficient. The decision was communicated during

a conference call involving top engineering managers, but it was not discussed

with the DER or the FAA. It was then documented in a memorandum that

contained no indication that it represented an extension of PS960A, and made no

reference to shotpeened taper bores. The memorandum stated only that blades

returned as a result of an ultrasonic inspection should be reworked “per

PS960A.”

The substance of the decision was then verbally transmitted by the

engineering manager of the Rock Hill facility to his staff but, as evidenced by the

technician’s belief that he was authorized to use the PS960A blend repair to

remove ultrasonic indications on both shotpeened and unshotpeened blades, it was

either misstated or misunderstood.

The blending process could mask the existence of a crack if done improperly (see

Section 2.2.1.2, discussing effects of failing to restore the original surface finish) or if enough of

the crack is removed by the blending so that the ultrasonic indication is reduced to a

nonrejectable height.

See Section 1.16.3 for more information about the circumstances surrounding the

decision to extend PS960A.

Although Hamilton Standard management asserted that this

expansion of the use of the PS960A blending repair procedure applied only to

ultrasonic indications in shotpeened taper bores, it was understood, at least by the

technician who worked on the accident blade, as being applicable to unexplained

ultrasonic indications in unshotpeened taper bores as well. Given that

unexplained ultrasonic indications in the taper bore area represent an unknown

condition suggestive of cracking and, further, that (according to statistical data

provided by Hamilton Standard) blades without shotpeened taper bores are

susceptible to earlier corrosion and to cracking once corrosion begins, Hamilton

Standard management (both in Windsor Locks and in Rock Hill) should have

made certain that the technicians performing the repair clearly understood that the

extension of PS960A was intended for shotpeened taper bores only.

If the technician had clearly understood that he was not authorized to

blend unexplained ultrasonic indications in unshotpeened taper bores, he would

have rejected the accident blade, or at least sought additional guidance from his

engineering manager as to how to handle the unexplained ultrasonic indication. In

either case, the accident blade would not have been subjected to the PS960A blend

repair, which masked the existence of the crack (see Section 2.2.1.2), and would

not likely have been returned to service. The Safety Board concludes that the

manner in which the unapproved extension of PS960A was documented and

communicated within Hamilton Standard, and the lack of training on the

extension, created confusion and led to misapplication of the blending repair to

unshotpeened blades with unexplained ultrasonic indications, allowing the

accident blade to be placed back into service with an existing crack, thus

contributing to this accident. The Safety Board is concerned about the adequacy

of Hamilton Standard’s internal communication and documentation. This issue is

further addressed in Section 2.6.1.

2.2.1.2 Sanding (Blending) of the Accident Blade

As noted in Section 1.16.1, although PS960A required that the

blended area be restored to its original surface finish, the sanding marks in the

blended area of the accident blade were much rougher than the original surface

finish. The sanding marks left by the blending appeared to have smeared some of

the corroded surface, suggesting that the sanding took place after the corrosion

had formed. Although some of the fatigue initiation area was along the sanding

marks, the fatigue cracking initiated from corrosion pitting damage that extended

below the taper bore surface to a depth much greater than the sanding marks.

Therefore, the Safety Board concludes that the sanding marks left by the PS960A

blending repair did not contribute to the initiation of the fatigue crack in the

accident blade. However, as discussed below, the sanding marks may well have

allowed the cracked blade to pass the ultrasonic inspection following the PS960A

blending repair.

Following the PS960A blending repair, the accident blade was again

ultrasonically inspected, and this time, neither a rejectable

nor a reportable

indication was generated. This allowed the blade to be returned to service, and it

was installed on the accident airplane. Safety Board laboratory measurements

indicated that only a minimal amount of material (less than about 0.002 inch) was

removed from the taper bore surface during the PS960A blending repair process.

Although removal of this amount of material would have slightly decreased the

size of the crack, it is unlikely that such a small decrease, by itself, would have

reduced the ultrasonic indication from 60 or 52 percent (the magnitude of the

indications recorded in the previous on-wing inspection and upon receipt at Rock

Hill, respectively) to below 40 percent (the minimum level that must be recorded

on the inspection form).

The ultrasonic inspection beam must reflect off both the defect and

the taper bore surface to bounce back to the transducer and be detected. However,

the surface finish in the taper bore was much rougher than the original finish, and

the sanding scratches that remained in the taper bore surface would have caused

the ultrasonic beam to scatter or become diffused as the beam reflected from the

taper bore surface. It is likely that this was the major reason for the decrease in the

size of the ultrasonic indication. Therefore, the Safety Board concludes that the

failure to restore the taper bore surface to the original surface finish during the

blend repair, as required by PS960A, was a factor that caused the reduction of the

ultrasonic indication that allowed the blade to pass the final ultrasonic inspection

and to be returned to service, and thus contributed to the accident.

In trying to determine why the technician did not properly perform

the blend repair, the Safety Board noted that PS960A did not describe or refer to

procedures that would ensure that the blended area was restored to its original

surface finish but simply stated this as a requirement. The technicians were

provided with examples of the required surface finish, which were designed to

50 percent or over of full-scale height.

40 percent or over of full-scale height.

assist them in determining whether the blended surface finish was appropriate.

However, Safety Board investigators found that it was difficult to distinguish

whether a particular surface finish varied from the examples. Investigators also

considered the possibility that the technician’s improper performance of the blend

repair was related to inadequate training in the repair,

his automotive

background, or the long hours he was working at the time.

However, after

extensive records examination and questioning of the technician, his supervisor

and others at Hamilton Standard, it was not possible to determine whether, or to

what extent, these factors contributed to the technician’s failure to properly restore

the blended area to its original surface finish.

Although the PS960A blend repair is no longer being used, Hamilton

Standard uses blending (sanding) in a variety of other propeller repair procedures.

In view of the potential for improperly performed blend repairs to mask existing

corrosion and cracks, the Safety Board believes that the FAA should require

Hamilton Standard to review and evaluate the adequacy of its tools, training and

procedures for performing propeller blend repairs, and ensure that those blend

repairs are being performed properly.

The technician who performed the blend repair on the accident blade

was neither an FAA-certificated mechanic nor, as an employee of a 14 CFR Part

145 repair station, was he required to be certificated. The technician stated that he

was permitted to sign off the work that he was qualified to perform. The shop

traveler form, which listed the requirement for the 63 RMS surface finish, showed

that the technician had signed off that he had accomplished the taper bore blend

repair. However, except for the subsequent ultrasonic inspection that was to

determine if the rejectable indication had been eliminated, there were no other

inspections of the accident blade. 14 CFR Part 65.87 states, in part, that a

certificated mechanic may return a propeller blade to service after he has repaired

and inspected, or supervised the repair and inspection of, that part. 14 CFR Part

145.45 specifies that a repair station must have an inspection system with qualified

personnel to determine the airworthiness of the parts being altered or maintained.

Therefore, the Safety Board believes that the FAA should review the need to

require inspection (“buy back”) after the completion of work that is performed by

He received 90 hours of on-the-job training from the engineering manager, whose

primary duties did not relate to training.

His work records indicate that during May and June of 1994, he worked between 8 and

26 hours of overtime each week.

uncertificated mechanics at Part 145 repair stations to ensure the satisfactory

completion of the assigned tasks.

2.2.2 Adequacy of Hamilton Standard Procedures for Detecting

Corrosion

2.2.2.1 Borescope Inspection

From an engineering and airworthiness perspective, the

appropriateness of the PS960A blending repair procedure rests on the premise that

there are no cracks or corrosion in the area being blended. However the

investigation revealed that during the initial stages of inspecting and repairing the

490 returned blades, although great emphasis was given to the development and

implementation of the PS960A repair procedure, little consideration was given to

the adequacy of the procedures being used to detect cracks and corrosion.

The inside of the taper bore is less than 1 inch in diameter. Visual

inspection of the interior of the taper bores was accomplished using a borescope

that did not provide low angle (indirect) illumination, which would have

highlighted changes in the surface depth, such as those related to subtle corrosion.

Instead, the borescope inspection used direct illumination, which produced a glare

and made visual detection of corrosion difficult. Using this direct-light borescope

resulted in a view of the taper bore interior that would reveal gross discoloration

or widespread corrosion but that would not readily reveal minor corrosion or small

cracks. In addition, the glare could have caused eye strain over time. For these

reasons, the probability of detection of cracks or corrosion decreased over the

course of a work period.

Corrosion in earlier blade failures, such as the Inter-Canadien

accident, was widespread and deep, and most likely would have been readily

detected by a visual inspection, even with a direct light source. However, the

corrosion involved in the accident blade was shallow, diffused and more difficult

to recognize as corrosion. In fact, the ultrasonic inspection performed prior to the

PS960A repair directed the technician to the location of the corrosion pits, but

upon visual examination of that area, the technician reported that he did not see

“visible faults.” Safety Board investigators, and other experts who have examined

the accident blade, agreed that inspection with a direct light source might not have

adequately highlighted the corrosion and cracking that had already developed.

(An FPI inspection technique was later introduced, using a black light borescope,

which would more readily reveal cracks and corrosion.)

In addition, at the time the accident blade was inspected, the

movement of the borescope into the taper bore hole and circumferentially around

the hole was controlled by hand, and there was no procedure to ensure that the

entire inner surface was inspected. Therefore, it is possible that the visual

inspection of some blades, possibly even the accident blade, might have been

incomplete. (Mechanical assistance to control the location of the borescope was

introduced later.)

Therefore, the Safety Board concludes that the borescope inspection

procedure developed and used by Hamilton Standard in June 1994 to inspect

returned blades that had rejectable ultrasonic indications for evidence of cracks,

pits, and corrosion was inadequate and ineffective.

2.2.2.2 Technician Training and Supervision

Investigators were informed that technicians at Hamilton Standard,

Rock Hill, normally received approximately 250 hours of general training before

they began working on propellers being returned to airworthy condition. Because

the technician who accomplished the taper bore repair to the accident blade

transferred from another shop area (fiberglass and nickel sheath replacement and

repair), he received his 250 hours of training in that area. After transferring to the

taper bore area, he received about 90 hours of on-the-job training (OJT) for taper

bore repair. However, that training was deficient because he was not provided a

photograph or model of what corrosion or cracking looked like inside a taper bore.

It was not until the fall of 1994 that such material was provided (in the form of a

photographic example of corrosion and an instructional video), as part of a group

of inspection improvements covered in a Component Maintenance Manual

revision and Service Bulletin 14RF-9-61-70A. Compliance with that Service

Bulletin was not required until March 23, 1995, with AD-95-05-03.

In the absence of a photograph or model of different types of potential

corrosion, the technician might have expected the type of gross corrosion or

cracking more familiar to his automotive background and, thus, failed to realize

the significance of corrosion with a more subtle appearance, such as that present in

the accident blade. This problem was most likely compounded by the poor

borescope inspection techniques discussed above. Further, technician training in

how to perform the taper bore repair procedure was provided by the Rock Hill

facility engineering manager, who had numerous additional responsibilities at the

time. Moreover, because the other inspectors were also recently hired and trained

and were relatively inexperienced in finding corrosion in the taper bore of

aluminum aircraft propellers, they were unable to provide guidance or help to each

other.

The Safety Board concludes that although the introductory technical

training to prepare the new, inexperienced workforce at Hamilton Standard’s Rock

Hill Customer Service Center may have been adequate, the specific training

initially given to technicians who inspected blades returned to Rock Hill as a

result of the on-wing ultrasonic inspections, including the accident blade, was not

adequate to ensure that they were proficient in the detection of taper bore

corrosion or associated cracks. (Hamilton Standard eventually improved its

training by providing photographs and an instructional video.)

2.2.2.3 Adequacy of Improvements to Inspection Procedures

The Safety Board is concerned that Hamilton Standard and the FAA

determined that repetitive ultrasonic inspections (per AD 95-05-03) could be

terminated by a one-time visual borescope inspection to detect corrosion or by a

previous PS960A repair, even though the probability of detection (POD

) for the

visual borescope method would be less than 100 percent. As a result of this

determination, when the ASA accident blade was visually inspected, repaired, and

marked with “PS960A” (which incorporated a borescope inspection of the taper

bore), the blade was exempt from further ultrasonic or visual inspections.

Hamilton Standard’s and FAA’s expectations for a visual inspection were

unrealistic, particularly when using hand-held borescopes with a direct light

source, and when the technician did not have explicit examples of what was a

rejectable visual condition.

In the months following the first two failures in March 1994,

Hamilton Standard identified inadequacies in the inspection process and generated

improvements to address these inadequacies. Specifically, Hamilton Standard

provided a photographic example of corrosion, a training video, a borescope feed

adapter, and a process upgrade to FPI. However, when these improvements in the

The probability of detection (POD) for a visual inspection is dependent upon many

variables including the flaw size, environment, ocular assistance devices, training and operator

experience and fatigue.

inspection methods were made, Hamilton Standard either did not recognize or was

not concerned that taper bore flaws, such as the crack in the ASA blade, might

have gone undetected during the previous inspection and repair process before the

improvements were made because Hamilton Standard did not implement

retroactive inspection of those blades that had been inspected previously and

returned to service under inspection standards and processes that were no longer

considered adequate.

Because the accident blade had been rejected once for an ultrasonic

indication caused by the crack that ultimately propagated to cause separation, it is

very likely that the existing crack would have been redetected if the accident blade

had been subjected to additional ultrasonic inspections (at the required 1,250 cycle

intervals, per AD 95-05-03) after it was returned to service following the Rock

Hill maintenance activities. If the accident blade had been subjected to the

recurrent on-wing ultrasonic inspections required by AD-95-05-03, it would have

received one inspection, and possibly a second, before the crack reached critical

size and the blade separated. Therefore, the Safety Board concludes that Hamilton

Standard’s failure to recommend, and the FAA’s failure to require, repetitive

ultrasonic inspections for all propellers (particularly those already inspected when

there were recognized shortcomings in the inspection process) contributed to the

accident because the crack in the accident blade would likely have been detected

in a recurrent ultrasonic inspection.

2.3 Effect of Blade Resonance

When Hamilton Standard attempted to model the growth of a taper

bore fatigue crack in the fall of 1995, the results of the initial fracture mechanics

analysis of the fracture of the Inter-Canadien (ATR-42) blade were consistent with

the initiation and propagation of the crack from the corrosion pits in that blade.

However, the analysis indicated that cracking should not have developed from the

smaller corrosion pits that were found at the origin area of the ASA blade and the

Nordeste blade (both from EMB 120 airplanes).

To more fully understand the ASA taper bore fracture mechanics,

Hamilton Standard then used a NASA-developed program called FASTRAN,

which addressed the magnitudes and numbers of stress cycles required for

corrosion pits to form cracks and for the cracks to propagate to critical size. Using

the measured flaw sizes from the accident blade and the two previous fractures,

the FASTRAN program was able to predict the stress cycles needed for the crack

in each of the blades, including the ASA blade, to propagate from the initial flaw

size to fracture. Hamilton Standard indicated to the Safety Board that for a

corrosion pit to initiate a crack, especially for small pit sizes, such as in the ASA

blade, a blade would have to have been subjected to a very high number of stress

cycles of the most severe type that the blade would normally encounter in routine

operations. Since a ground-air-ground (GAG) cycle

imparts severe stress to the

blade only once per flight, Hamilton Standard engineers believed that 2P

resonance (which occurs twice per revolution of the propeller) in adverse winds,

perhaps during a maintenance ground run, contributed to the initiation of the crack

from corrosion pitting and propagation of the crack while it was small. The

FASTRAN program calculated that for a crack the size of the ASA crack at the

time the blade was at Rock Hill (0.050 by 0.060 inch) to propagate to failure, an

average of 50 maximum level 2P stress cycles would have to be accumulated each

flight. This number of cycles could be accumulated in about 1.6 seconds of

operation at the rpm range associated with 2P resonance during ground operation

in adverse winds (quartering tailwinds).

Through independent testing using a corporate-owned EMB-120

airplane, Hamilton Standard confirmed previous test results that the most severe

stress cycles were the GAG cycles, and the vibratory stresses encountered during

2P resonance (which are about the same magnitude as the GAG stresses) in

adverse winds during ground operations.

It should be noted that during the propeller ground tests conducted by

Hamilton Standard (where stress levels up to 18,000 psi were measured), the test

lasted for only 30 seconds. The 30-second interval might not have been of

sufficient duration, with the given test conditions, to produce maximum possible

stress. Additionally, the quartering tailwinds were generated by the propeller

wash of a small executive turbopropeller aircraft. It is possible that stress levels in

excess of 18,000 psi could be attained if the propeller were allowed to remain in

the resonant condition in excess of 30 seconds or if the propeller was subjected to

natural quartering tailwind conditions of higher velocity.

A GAG cycle included engine power application for takeoff, climb, cruise, descent and

reverse thrust after landing.

Ground operation in a tailwind or quartering tailwind causes the airplane to shake and

buffet, which hampers the mechanic’s ability to read instruments and perform any meaningful

tests.

The Safety Board noted that ASA’s operational procedures prohibited

single-engine taxi. Because single-engine taxi increases the potential for ground

operation in the resonant range, by requiring higher engine rpm to initiate and

continue movements, it might be expected that other operators of EMB-120

airplanes that allow single-engine ground operation would have a higher incidence

of cracks. That is, if a 2P resonant condition could be generated during normal

ground operations and could initiate and propagate a fatigue crack, it would be

more likely that cracking would occur on blades used by operators that allow

single engine operation on the ground. However, this did not occur.

It is also possible that a 2P resonant condition was generated during

some type of ground maintenance activity and that these stresses contributed to the

initiation and propagation of the crack. However, ground operation in a tailwind

or quartering tailwind is contrary to all standard maintenance practice and is

prohibited in the EMB-120 maintenance manual. The Safety Board could find no

documented evidence, and was not provided any information by Hamilton

Standard, Embraer, Pratt and Whitney of Canada, or the operator, that the

propeller was operated in a 2P resonant condition. Additionally, the EMB-120

cockpit was placarded, and there were cautions set forth in the airplane flight

manual and engine maintenance manual to avoid engine operation above 65

percent Np range in adverse wind conditions. Safety Board investigators had a

number of opportunities to observe whether there was compliance with these

limitations, and they did not observe noncompliance.

Although ground operation in a tailwind causes the airplane to shake,

the shaking is a result of disturbed airflow across the wings and fuselage and is not

damaging. However, under such conditions, an increase of the propeller rpm into

the resonant range would generate a reactionless vibratory mode in the propellers.

This would be undetectable, especially during crosswind-induced airframe buffet.

Such circumstances could arise at any time during ground operation in a strong

quartering tailwind, such as while operating in the vicinity of jet wash from a large

transport aircraft. With a population of more than 2,450 14RF-9 model blades in

service on EMB-120 airplanes throughout the world, there have probably been

many occasions when blades were operated in quartering tailwinds in excess of the

10 knots cautioned about in the flight manual and placards.

The Safety Board concludes that a combination of 2P resonance and

GAG cycle stresses initiated the crack from the corrosion pits in the ASA blade

and caused the crack to propagate to failure under normal operating conditions.

2.4 Adequacy of Vibration Testing

Hamilton Standard’s 1983 and 1985 certification stress surveys were

conducted using one instrumented blade with a second instrumented blade as a

backup. None of the tests were conducted using blades that had previously been

used in service. The 1995 stress survey conducted as a consequence of this

accident included three instrumented blades that had approximately 12,000 service

hours. The 1995 stress survey revealed a small but important difference from the

1983 and 1985 certification tests. Although no new high stress conditions were

discovered, a small downward shift in the first flat-wise natural frequency was

detected. This downward shift was attributed to normal wear and mass properties

in the propeller blade from normal operation.

On April 19, 1993, the Safety Board investigated a propeller in-flight

separation on an MU-2B-60 in Zwingle, Iowa. As indicated in the accident report,

the Safety Board discovered that during certification testing of the Hartzell HC-B4

propeller on the MU-2B airplane, a reactionless mode of vibration was identified

with the peak stress occurring at a propeller speed of 1,079 rpm. As a result, the

propeller was prohibited from continuous operation on the ground below 1,145

rpm. The Safety Board did not find evidence that the test was repeated using

propeller blades altered to conform to the minimum dimensions specified in the

repair limit criteria contained in the Hartzell HC-B4 propeller maintenance

manual.

The FAA recommends in Advisory Circular (AC) 20-66, Vibration

Evaluation of Aircraft Propellers, in the chapter on vibration measurement

programs, that propeller diameters should be tested at various lengths throughout

the diameter range, to include the maximum and minimum diameters, and the

cutoff repair limit. Although it is not explained in AC 20-66, testing a propeller

blade at its cutoff repair limit is conducted presumably to ensure that the decrease

in blade length does not alter the natural frequency to the extent that a resonant

vibration condition could be entered while operating within the normal rpm range.

AC 20-66 considers the expected loss of mass of a metal propeller blade following

blend repair; however, the AC does not consider the expected mass gain with age

of a composite propeller blade. Composite blades may gain mass with age with

the addition of layers of paint, introduction of moisture, and patch repairs.

A review of AC 20-66, which includes a detailed discussion of the

propeller vibratory phenomenon, does not explain that a propeller blade’s natural

vibratory response varies with varying conditions (i.e. mass gain, mass loss,

variations in airfoil shape, etc.) and that adequate margin from a potentially

coincident excitation frequency should be maintained. Consequently, the Safety

Board concludes that the AC does not provide guidelines for adequate margin

between a propeller blade’s natural frequencies and its potentially coincident

excitation frequencies over the life of the blade.

Therefore, the Safety Board believes that the FAA should revise AC

20-66 to include the vibratory testing of composite propeller blades that have been

previously operated for a substantial number of service hours, and composite

blades that have been altered to the limits set forth in FAA-approved repair

manuals to determine the expected effects of age on propeller vibration and

provide guidelines for rpm margin between a propeller blade’s natural frequencies

and the excitation frequencies associated with propeller operation.

2.5 Effect of Blade Failure and Analysis of Terminating Action

The forward half of the fractured front inlet case, found along the

wreckage path, was attached to the reduction gear box (RGB) of the left engine.

The RGB case was also fractured, empty of oil, and some internal gearbox

components were missing. The in-flight fracture of the RGB case most likely

resulted in the loss of the gearbox components and the venting liquid reportedly

seen by the passengers.

Damage to the forward engine mounts and the front frame of the left

engine nacelle indicated that the propeller RGB first separated upward at the

inboard mount location, followed by the aft and outboard failure at the outboard

mount location. Failure of the engine mounts were the result of overload stresses.

The imbalance loads associated with the separated blade on the

accident airplane apparently exceeded the design strength of the RGB attachment

points. In this instance, it appears that the RGB was retained on the airplane in an

outboard displaced position, as observed by the passengers, from the time of

propeller blade separation until the beginning of the initial ground impact.

Embraer’s postcertification (September 1984) testing determined that

the nacelle would not withstand a mid-blade or full-blade segment loss. To date,

there have been four blade separations--three from fatigue cracks that initiated in

the taper bore. The first blade separation (Inter-Canadien) resulted in RGB and

propeller separation, and the assembly fell to earth. During the second blade

separation (Nordeste), the RGB and propeller assembly remained in place. During

the third separation (the Luxair accident), in which a fracture occurred in the blade

shank area, the RGB and propeller assembly again fell from the airplane. During

the fourth blade failure (this accident), the RGB rotated out of position, and

resulted in degraded aerodynamic performance and a fatal accident.

Although in two of the occurrences, the RGB and propeller fell clear

and did not seriously compromise the airplane or degrade its performance, all of

the occurrences clearly placed the airplane and its occupants at potential serious

risk. On four occasions, stresses on blades with flaws (corrosion pits or

mechanical damage) have produced a blade separation even though the propeller

was certificated based on the assumption of an unlimited life. Because the current

regulations do not require that an airframe survive if a blade breaks, and because

Embraer has determined that the EMB-120 cannot survive the loss of a mid-blade

or full-blade segment, minimizing the possibility of a propeller blade separation is

imperative. To prevent future failures, it is essential that stress risers in the form

of corrosion or mechanical damage are not permitted to occur on any propeller

blade.

The Safety Board concurs that the taper bore repair procedure

specified in the March 1996 Service Bulletins (and required by ADs) should have

restored the surface of the taper bore of all existing propellers to a nearly new

condition. Also, because Hamilton Standard has prohibited the use of the

mechanical lead-removal tools during routine blade balancing, the likelihood of

future inadvertent mechanical damage has been greatly reduced.

However, while the terminating taper bore repair procedure should

detect and eliminate any chlorine-induced corrosion or mechanical damage, the

Safety Board is concerned that exposure to small amounts of moisture or other

atmospheric elements during routine maintenance, the recurring inspection

procedure set forth in the CMM, periods of low utilization, or long-term storage

may allow atmospheric-induced corrosion to begin in the taper bore. The Safety

Board is aware of reports of corrosion and cracking in the taper bores of P-3 and

C-130 propellers associated with long-term storage. Because of this, the Safety

Board concludes that despite all the actions taken by Hamilton Standard and the

FAA to date, there is a continuing potential for corrosion to develop in taper bores

of the affected Hamilton Standard propeller blades. Therefore, the Safety Board

believes that the FAA should require that Hamilton Standard consider long-term,

atmospheric-induced corrosion effects and amend the CMM inspection procedure

to reflect an appropriate interval that will detect any corrosion within the taper

bore.

2.6 FAA Oversight

2.6.1 Role of Designated Engineering Representative and FAA

Certifying Engineer

As discussed in Section 1.16.3, Hamilton Standard extended the

applicability of the PS960A blending repair beyond its original purpose (blending

of mechanical damage) to also allow blending of shotpeen impressions to remove

ultrasonic indications without informing or involving its DER or the FAA

(through its Certifying Engineer in the Engine and Propeller Directorate).

Hamilton Standard was aware that the FAA Certifying Engineer had earlier

determined that repair procedures to the taper bore would require direct FAA

approval. Consistent with this determination, both the DER (who served as a

liaison to the FAA) and the Certifying Engineer were involved in developing the

initial ultrasonic inspection procedures later set forth in ASB 14RF-0-61-A66 and

the PS960A repair procedure. If Hamilton Standard had consulted with the DER

and the FAA about the proposed extension and the reason for it (ultrasonic

indications in blades without any visible damage), it might have prompted the

FAA to reconsider whether the inspection and repair processes then being used

were adequate to identify blades at risk for cracking.

Even though the extension of PS960A to allow blending of shotpeen

impressions might have been approved by the FAA, if approval had been

requested, Hamilton Standard’s failure to seek such approval nonetheless played a

role in creating the confusion that led to the misapplication of PS960A. The

FAA’s approval would likely have been documented in a fashion similar to its

approval of PS960 and PS960A, clearly specifying the limits of the approved

extension. In contrast, as discussed in Section 2.2.1.1, Hamilton Standard’s

attempt to internally document the extension was confusing and led to the

misapplication of the repair to unshotpeened taper bores with unexplained

ultrasonic indications.

Accordingly, the Safety Board concludes that Hamilton Standard’s

failure to seek FAA approval of the extension of PS960A blending repair hindered

the FAA’s ability to oversee Hamilton Standard’s handling of the taper bore crack

and corrosion problem, and led to an inadequate documentation of the extension

that caused confusion and misapplication of the repair. Although the DER stated

that this lapse in communication was atypical, the Safety Board is concerned--

especially in light of the inadequate manner in which Hamilton Standard

communicated the information to its managers and technicians--that it may

represent a deficiency in Hamilton Standard’s corporate communication.

Specifically, it suggests that Hamilton Standard placed insufficient emphasis on

proper communication of vital safety information. Accordingly, the Safety Board

believes that the FAA should require Hamilton Standard to review and, if

necessary, revise its policies and procedures regarding 1) internal communication

and documentation of engineering decisions, and 2) involvement of the DER and

FAA, and to ensure that there is proper communication, both internally and with

the FAA, regarding all significant engineering decisions.

2.7 Weather

The flight was operating in accordance with an IFR flight plan and

was climbing above overcast clouds at the initiation of the accident sequence. A

weather observation taken at CTJ, approximately 4 miles from the crash site,

reported an 800-foot overcast cloud ceiling just after the accident. However,

weather conditions appropriate for visual flight below the clouds were reported at

several airports near the accident site.

From the flightcrew’s requests to ATC for vectors to the airfield, it is

apparent that the cloud ceiling affected the flightcrew’s ability to visually acquire

a suitable landing site during the descent for a forced landing. In the latter portion

of the descent, after descending below the overcast cloud ceiling, the airplane’s

height above the terrain would have limited the view of the flightcrew to just the

immediate area. The airplane impacted the ground in a left-wing-down attitude,

probably because the flightcrew was attempting to complete a turn to properly

align themselves for the forced landing. If the overcast cloud ceiling had been

higher, the crew would have had more time to align the airplane and level the

wings before the impact. Consequently, the Safety Board concludes that the cloud

ceiling precluded the flightcrew from being able to see the ground and thus to

make a more successful forced landing, a situation that contributed to the severity

of the accident.

2.8 Air Traffic Control Services

All FAA ATC personnel were trained, certificated, and qualified for

their duties. There were no apparent physiological or behavioral impairments or

disabilities that would have detracted from their ability to provide the expected air

traffic services. Following the declaration of the emergency by the flightcrew and

their continuing description of their difficulties, the controllers provided

appropriate assistance to the crew.

The Atlanta Center controller did not issue the CTJ AWOS

frequency. Although FAA ATC procedures

require that “both center and

approach controllers shall provide current approach information to aircraft

destined to airports for which they provide approach control services,” in this case,

the Atlanta Center controller was not responsible for approaches into CTJ and was

not in the best position to have provided this information.

The closest weather report immediately available to the approach

controller was the ATL Airport observation, the flight’s departure point. There

was no controller assigned to the “assist” position and although the manager and

supervisor were nearby, they became occupied with coordinating and monitoring

activities supporting the flight and did not attempt to retrieve the CTJ AWOS

weather information by telephone. During the 90 seconds that the approach

controller was in radio communication with the flight, the controller issued a

vector toward the runway, the localizer frequency, confirmed the flight was in

visual conditions and issued a vector for the visual approach. Therefore, the

Safety Board concludes that, although the Atlanta approach controller did not

issue the AWOS frequency or provide weather information, the controller

performed higher priority tasks and, because the flight had to land at the nearest

airport regardless of the weather, the failure to provide the CTJ weather

information to the flightcrew was not a factor in this accident.

The Safety Board examined the performance of both the center and

the approach controllers in assisting the accident flight. Once the flightcrew

declared the emergency, Atlanta Center immediately issued instructions to permit

the flight to return to ATL. Because of the location of the aircraft within the

sector, the radar controller and manual assist controller coordinated with

FAA Order 7110.65, “Air Traffic Control,” Chapter 4, “IFR,” Section 7, “Arrival

Procedures,” paragraph 4-7-13, “Approach Information,” applies.

surrounding sectors; the center controller attempted to assist the accident

flightcrew by providing vectors and airport information. However, because of the

degraded flight performance of the aircraft, the flightcrew was not able to effect a

timely response to the controller’s instructions. When the flightcrew requested

information normally available from the approach controller, such as runway

heading and airport surface condition, the center controller instructed the

flightcrew to contact the approach control facility. FAA ATC procedures

state in

part, “If you are in communication with an aircraft in distress, handle the

emergency and coordinate and direct the activities of assisting facilities. Transfer

this responsibility to another facility only when you feel better handling of the

emergency will result.”

The flightcrew would have been able to advise the center controller

that they preferred to remain on the current frequency because of the emergency,

and the center controller may have been able to coordinate alternate procedures.

Also, in retrospect, the Atlanta Center controller could have made an earlier

handoff if he had been informed by the flightcrew of, or if he had perceived the

full extent of, the performance degradation to the accident airplane. However, the

Safety Board concludes that the timing of the handoff to Atlanta approach control

by the Atlanta Center controller was not a factor in the accident.

The Safety Board is concerned, however, about the failure of ATC

controllers to notify CFR services once the controllers were aware of the

emergency situation. At 1644:25, the flightcrew of ASE 529 notified the Atlanta

Center air traffic controllers that they had experienced an engine failure and

declared an emergency. Two minutes later, the flightcrew advised that they

needed to “land quick” and requested the controller to “roll the trucks and

everything for us.” The controller then advised the flightcrew that CTJ was the

closest airport and directed the aircraft to CTJ. Although ATC was aware of the

emergency situation and destination airport, ATC did not notify the fire and

emergency services covering CTJ, the Carroll County Fire Department, of the

incoming aircraft.

Atlanta Center should have immediately advised the appropriate CFR

service or instructed Atlanta approach of the pilot’s request so that they could have

made timely airport emergency services notification. The accident had already

FAA Order 7110.65, Chapter 10, “Emergencies,” Section 1, “General,” paragraph 10-1-

4, “Responsibility,” applies.

occurred when the Atlanta approach controller made the call to the Carroll County

Sheriff, and it had already been reported by a citizen on 911. The Safety Board

concludes that if the Atlanta Center had placed a call for emergency services as

soon as the pilot requested, which was 10 minutes before the accident, personnel

would have responded sooner, and the rescue efforts might have been more timely

and effective. Therefore, the Safety Board believes that the FAA should include

an article in the Air Traffic Bulletin and provide a mandatory formal briefing to all

air traffic controllers regarding the necessity and importance of notifying crash,

fire and rescue personnel upon a pilot’s request for emergency assistance. Ensure

that air route traffic control center (ARTCC) controllers are aware that such a

request may require them to notify local emergency personnel.

2.9 Survival Factors Aspects

The Safety Board commends the exemplary manner in which the

flight attendant briefed the passengers and handled the emergency. According to

passengers, immediately following the loss of the propeller blade, the flight

attendant checked with each passenger individually to make sure that they all

understood how to assume the brace position, and she yelled instructions to the

passengers up until the time of impact. After the crash, although she was seriously

injured, she continued to assist the passengers by moving them away from the

airplane and extinguishing flames on at least one passenger who was on fire.

2.9.1 Time Management During Emergencies

The Safety Board recognizes that the flightcrew in this accident was

attempting to control the aircraft. However, the Safety Board is concerned that the

flight attendant neither received nor sought information about the time remaining

to prepare the cabin or to brace for impact. The CVR transcript revealed that the

flightcrew informed her 7 minutes before impact that they had experienced an

engine failure, that they had declared an emergency for return to ATL, and that

they had advised her to brief the passengers. There were no further

communications to the flight attendant. Specifically, the flight attendant was

never told that the airplane would not be able to make ATL, and would instead be

making an off-airport crash landing. The flight attendant stated that while

preparing the cabin and passengers, she saw the tree tops from a cabin window.

She immediately returned to her jump seat and shouted her commands. A

passenger commented that the flight attendant was barely in the brace position

when the impact occurred.

The Safety Board is concerned that the flight attendant and the

flightcrew did not discuss a brace signal and the time available to prepare the

cabin, and that the flightcrew did not announce a brace command on the public

address system. Further, if the flight attendant had not had sufficient time to

fasten her safety belt and shoulder harness, she might have received more serious

or fatal injuries, and she might have been incapable of directing an evacuation.

The FAA has recognized that communication and coordination

between cockpit crewmembers and flight attendants continue to challenge air

carriers and the FAA. Advisory Circular (AC) 120-51B, “Crew Resource

Management Training,” suggests several methods of addressing this problem.

Paragraph 15, Evolving Concepts of CRM: Extending Training Beyond the

Cockpit, addresses specific subjects for joint training but does not specifically deal

with the communication of critical information during an emergency.

The Safety Board’s special investigation report on flight attendant

training

describes another accident on page 28:

The lead flight attendant in the DC-10 stated that she knew

emergency procedures required her to determine the amount of

time available to prepare the passengers and the cabin. However,

she chose not to ask the flightcrew about the time. Additionally,

the second item on the flight attendant checklist was “Determine

Time,” but none of the flight attendants followed this checklist

procedure.

Although the FAA issued Air Carrier Operations Bulletin 1-91-11 in

response to Safety Board Safety Recommendation A-90-173, which called for

inspectors to reiterate the importance of time management in the preparation of the

cabin in a planned emergency, the Safety Board concludes that this accident also

illustrates that critical information regarding time available to prepare the aircraft

for an emergency landing or impact is not being considered and communicated

among flight and cabin crewmembers. Therefore, to improve the interactions

between the cockpit and cabin crews, the Safety Board believes that the FAA

should amend AC 120-51B to include guidance regarding the communication of

See Special Investigation Report—“Flight Attendant Training and Performance During

Emergency Situations” (NTSB/SIR-92/02)

time management information among flight and cabin crewmembers during an

emergency.

2.9.2 Crash Axes

The captain and first officer were trapped in the cockpit by fire that

had ignited on the cabin side of the cockpit door. When the first officer found it

impossible to open his cockpit sliding window, he unsuccessfully attempted to

chop a hole in the hardened Plexiglas side window using the airplane crash ax. It

was apparently intended for use as a woodworking tool because it consisted of a

blade and nail puller attached to a wooden handle. Given the resilient composition

of the cockpit window material, it was difficult to make a hole in the window

panel; however, if the ax had been equipped with a pry bar rather than a nail

puller, the first officer might have been successful in wedging the pry bar between

the window and the track or frame and prying or forcing the window open.

Although regulations exist that require most passenger-carrying aircraft to be

equipped with a crash ax,

there is no FAA or other civil technical standard

regarding the design and use of crash axes. This accident demonstrates the

importance of an adequate crash ax design.

The crash ax carried aboard military transport aircraft conforms to a

special design. Large commercial transport airplanes manufactured in the United

States are equipped with crash axes of similar design. Additionally, firefighter

axes that have a wedge and pry bar tool features are in use by airport rescue and

fire fighting personnel and municipal emergency medical technicians. The Safety

Board concludes that there should be standards governing the design of crash axes

required to be carried aboard passenger-carrying aircraft. Therefore, the Safety

Board believes that the FAA should evaluate the necessary functions of the

aircraft crash ax, and provide a technical standard order or other specification for a

device that serves the functional requirements of such tools carried aboard aircraft.

See 14 CFR 91.513(e), 135.177(a)(2), and 121.309(e).

3. CONCLUSIONS

3.1 Findings

1. The flightcrew was trained, certificated and qualified to

conduct the flight, and the flight was conducted in accordance

with applicable Federal Aviation Regulations and company

requirements.

2. The flightcrew was in good health and held the proper FAA

medical certificates. There was no evidence that the

performance of any crewmember was impaired by alcohol,

drugs, or fatigue.

3. ASA maintained the airplane in accordance with applicable

Federal Aviation Regulations and company Operations

Specifications.

4. After the propeller blade separation, the combination of the

resulting loss of left engine thrust, increased drag from a

deformed engine nacelle and the three blades retained in the

propeller hub and added frontal drag from external sheet

metal damage degraded airplane performance preventing the

flightcrew from arresting the airplane’s descent or making

rapid changes in its direction of flight making a forced

landing necessary.

5. One of the four blades from the left engine propeller

separated in flight because a fatigue crack that originated

from multiple corrosion pits in the taper bore surface of the

blade spar propagated toward the outside of the blade, around

both sides of the taper bore, then reached critical size.

6. Because of the severely degraded aircraft performance

following the propeller blade separation, the flightcrew’s

actions were reasonable and appropriate during their attempts

to control and maneuver the airplane throughout the accident

sequence, and they were not a factor in this accident.

7. Hamilton Standard’s engineering decision to use the PS960A

blending repair to remove ultrasonic indications caused by a

shotpeened taper bore surface was technically reasonable.

8. The manner in which the unapproved extension of PS960A

was documented and communicated within Hamilton

Standard, and the lack of training on the extension, created

confusion and led to misapplication of the blending repair to

unshotpeened blades with unexplained ultrasonic indications,

allowing the accident blade to be placed back into service

with an existing crack.

9. The sanding marks left by the PS960A blending repair did not

contribute to the initiation of the fatigue crack in the accident

blade.

10. The failure to restore the taper bore surface to the original

surface finish, as required by PS960A, was a factor that

caused the reduction of the ultrasonic indication that allowed

the blade to pass the final ultrasonic inspection and to be

returned to service.

11. The borescope inspection procedure developed and used by

Hamilton Standard in June 1994 to inspect returned blades

that had rejectable ultrasonic indications for evidence of

cracks, pits, and corrosion was inadequate and ineffective.

12. The introductory technical training to prepare the new,

inexperienced workforce at Hamilton Standard’s Rock Hill

Customer Service Center might have been adequate; but the

training initially given to technicians, who inspected blades

that were returned to Rock Hill as a result of on-wing

ultrasonic inspections, including the accident blade, was

inadequate to ensure proficiency in the detection of taper bore

corrosion or associated cracks.

13. If Hamilton Standard had recommended, and the FAA had

required, repetitive ultrasonic inspections for all propellers

after shortcomings were recognized and improvements were

made in the inspection process (particularly those that had

already been inspected), the crack in the accident blade would

most likely have been detected.

14. A combination of 2P resonance and GAG cycle stresses

initiated the crack from the corrosion pits in the ASA blade

and caused the crack to propagate to failure under normal

operating conditions.

15. Advisory Circular AC 20-66 does not provide guidelines for

adequate margin between a propeller blade’s natural

frequencies and its potentially coincident excitation

frequencies over the life of the blade.

16. There is a potential for corrosion to develop in taper bores of

the affected Hamilton Standard propeller blades.

17. The cloud ceiling precluded the flightcrew from being able to

see the ground and thus to make a more successful forced

landing.

18. Hamilton Standard’s failure to seek FAA approval of the

extension of PS960A blending repair hindered the FAA’s

ability to oversee Hamilton Standard’s handling of the taper

bore crack and corrosion problem, and led to an inadequate

documentation of the extension that caused confusion and

misapplication of the repair.

19. The timing of the handoff to Atlanta approach control by the

Atlanta Center controller was not a factor in the accident.

20. Although the Atlanta approach controller did not issue the

AWOS frequency or provide weather information, the

controller performed higher priority tasks; and because the

flight had to land at the nearest airport regardless of the

weather, the failure to provide the CTJ weather information to

the flightcrew was not a factor in this accident.

21. If the Atlanta Center had placed a call for emergency services

as soon as the pilot requested, which was 10 minutes before

the accident, personnel would have responded sooner, and the

rescue efforts might have been more timely and therefore

more effective.

22. This accident illustrates that critical information regarding

time available to prepare the aircraft for an emergency

landing or impact is not being considered and communicated

among flight and cabin crewmembers.

23. There should be standards governing the design of crash axes

required to be carried aboard passenger-carrying aircraft.

3.2 Probable Cause

The National Transportation Safety Board determines that the

probable cause of this accident was the in-flight fatigue fracture and separation of

a propeller blade resulting in distortion of the left engine nacelle, causing

excessive drag, loss of wing lift, and reduced directional control of the airplane.

The fracture was caused by a fatigue crack from multiple corrosion pits that were

not discovered by Hamilton Standard because of inadequate and ineffective

corporate inspection and repair techniques, training, documentation, and

communications.

Contributing to the accident was Hamilton Standard’s and FAA’s

failure to require recurrent on-wing ultrasonic inspections for the affected

propellers.

Contributing to the severity of the accident was the overcast cloud

ceiling at the accident site.

4. RECOMMENDATIONS

As a result of the investigation of this accident, the National

Transportation Safety Board makes the following recommendations:

--to the Federal Aviation Administration:

Require Hamilton Standard to review and evaluate the adequacy of

its tools, training, and procedures for performing propeller blend

repairs, and ensure that those blend repairs are being performed

properly. (A-96-142)

Review the need to require inspection (“buy back”) after the

completion of work that is performed by uncertificated mechanics

at Part 145 repair stations to ensure the satisfactory completion of

the assigned tasks. (A-96-143)

Revise Advisory Circular 20-66 to include the vibratory testing of

composite propeller blades that have been previously operated for

a substantial number of service hours, and composite blades that

have been altered to the limits set forth in FAA-approved repair

manuals to determine the expected effects of age on propeller

vibration and provide guidelines for rpm margin between a

propeller blade’s natural frequencies and the excitation

frequencies associated with propeller operation. (A-96-144)

Require that Hamilton Standard consider long-term, atmospheric-

induced corrosion effects and amend the Component Maintenance

Manual (CMM) inspection procedure to reflect an appropriate

interval that will detect any corrosion within the taper bore. (A-

96-145)

Require Hamilton Standard to review and, if necessary, revise its

policies and procedures regarding 1) internal communication and

documentation of engineering decisions, and 2) involvement of the

Designated Engineering Representative (DER) and FAA, and to

ensure that there is proper communication, both internally and

with the FAA, regarding all significant engineering decisions. (A-

96-146)

Include an article in the Air Traffic Bulletin and provide a

mandatory formal briefing to all air traffic controllers regarding

the necessity and importance of notifying crash, fire and rescue

personnel upon a pilot’s request for emergency assistance. Ensure

that air route traffic control center (ARTCC) controllers are aware

that such a request may require them to notify local emergency

personnel. (A-96-147)

Amend Advisory Circular 120-51B (Crew Resource Management

Training) to include guidance regarding the communication of

time management information among flight and cabin

crewmembers during an emergency. (A-96-148)

Evaluate the necessary functions of the aircraft crash ax, and

provide a technical standard order or other specification for a

device that serves the functional requirements of such tools carried

aboard aircraft. (A-96-149)

BY THE NATIONAL TRANSPORTATION SAFETY BOARD

James E. Hall

Chairman

Robert T. Francis II

Vice Chairman

John Hammerschmidt

Member

John J. Goglia

Member

George W. Black

Member

November 26, 1996

this page intentionally left blank

5. APPENDIXES

APPENDIX A

INVESTIGATION AND HEARING

1. Investigation

The Safety Board was notified of the accident by the FAA

Communication Center about 1330 on August 21, 1995. An investigator was

immediately dispatched to the crash site from the Southeast Field Office in

Atlanta, and a full go-team was sent from Safety Board Headquarters in

Washington, D.C. The following investigative groups were formed: Operations,

Structures, Systems, Powerplants, Maintenance Records, Air Traffic Control,

Weather, Aircraft Performance, Survival Factors, Cockpit Voice Recorder, and

Flight Data Recorder. Member John Hammerschmidt accompanied the team to

Carrollton, Georgia. A Metallurgy Group was later formed at the Safety Board’s

Materials Laboratory to study evidence of the failed propeller.

In accordance with the provisions of the International Civil Aviation

Organization’s International Standards and Recommended Practices, Annex 13,

Aircraft Accident and Incident Investigation, the Center for the Investigation and

Prevention of Accidents, Brazil, (state of manufacture of the airplane) and the

Transportation Safety Board of Canada, (state of manufacture of the engines) were

notified of the accident, and each state sent an Accredited Representative, with

advisers to participate in the investigation. A draft of the final accident report was

provided to each Accredited Representative for comment. Their comments have

been incorporated into the final report.

Parties to the investigation included the Federal Aviation

Administration, Atlantic Southeast Airlines, Empresa Brasileira de Aeronautica S.

A. (Embraer Aircraft Corporation), the Pratt and Whitney, Canada, Division of

United Technologies Corporation, the Hamilton Standard Division of United

Technologies Corporation, the Air Line Pilots Association, the Association of

Flight Attendants, and the National Air Traffic Controller’s Association.

2. Public Hearing

No public hearing was conducted for this investigation.

APPENDIX B

COCKPIT VOICE RECORDER TRANSCRIPT

LEGEND

HOT

Crewmember hot microphone voice or sound source

HOT-M

Aircraft mechanical voice heard on all channels

RDO

Radio transmission from accident aircraft

CAM

Cockpit area microphone voice or sound source

INT

Transmissions over aircraft interphone system

TWRA

Radio transmission from Atlanta control tower

ATLD

Radio transmission from Atlanta departure control

CTR

Radio transmission from Atlanta center

ATLA

Radio transmission from Atlanta approach control

UNK

Radio transmission received from unidentified aircraft

Transmission made over aircraft public address system

-B

Sounds heard through both pilot’s hot microphone systems

-1

Voice identified as Pilot-in-Command (PIC)

-2

Voice identified as Co-Pilot

-3

Voice identified as female flight attendant

Voice unidentified

Unintelligible word

Non pertinent word

Expletive

Break in continuity

( )

Questionable insertion

[ ]

Editorial insertion

....

Pause

Note: Times are expressed in eastern daylight time (EDT).

Times shown in brackets { } are computer reference times measured from the beginning of the recording.

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

START of RECORDING

START of TRANSCRIPT

1222:39 {00:22}

TWRA

AC five twenty nine, turn left heading zero six zero runway eight

right. cleared for takeoff.

1222:45 {00:28}

RDO-2

zero six zero eight right, cleared for takeoff, AC five twenty

nine.

1222:48 {00:31}

HOT-1

takeoff check, below the line, I got your lights.

1222:52 {00:35}

HOT-2

condition levers? max.

1222:52 {00:35}

HOT-2

bleeds and packs? standard.

1222:53 {00:36}

HOT-2

external lights? on.

1222:54 {00:37}

HOT-2

below the line's complete.

1222:55 {00:38}

HOT-1

alright.

1223:00 {00:43}

CAM

[sound of increase in rpm similar to takeoff power being applied]

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1223:19 {01:02}

HOT-2

power set, auto-feather's armed, panel's clear, airspeed’s alive.

1223:24 {01:07}

HOT-2

V 1.

1223:25 {01:08}

HOT-2

V r.

1223:29 {01:12}

HOT-2

pos' rate.

1223:30 {01:13}

HOT-1

gear up.

1223:37 {01:20}

HOT-1

wipers off.

1223:46 {01:29}

TWRA

AC five twenty nine, contact departure. fly heading zero six

zero now. we'll see ya.

1223:49 {01:32}

RDO-2

zero six zero switching, see ya.

1223:51 {01:34}

HOT-B

[single beep similar to radio frequency change]

1224:00 {01:43}

RDO-2

Atlanta departure, AC five twenty nine's with you, leaving one

point eight for four, heading sixty.

1224:07 {01:50}

ATLD

AC five twenty nine, Atlanta departure roger, radar contact.

maintain one zero thousand.

1224:13 {01:56}

RDO-2

one zero thousand, AC five twenty nine.

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1224:17 {02:00}

HOT-2

**.

1224:17 {02:00}

HOT-1

ten.

1224:31 {02:14}

HOT-1

climb power, * takeoff.

1225:04 {02:47}

HOT-2

gear's up, flaps are up, pressure's checked, climb power set, bleeds

and packs auto, APU's off, auto-feather off, flight attendant notified

****.

1225:11 {02:54}

HOT-1

thank you.

1225:26 {03:09}

RDO-2

five twenty nine's up, at twenty four.

1225:27 {03:10}

HOT-B

[single beep similar to radio frequency change]

1225:41 {03:24}

HOT-1

is this the s*, this the ## that uh, kept doin' that to us? on takeoff?

1225:54 {03:37}

HOT-2

yeah this is the one, that these were loose.

1225:57 {03:40}

ATLD

AC five twenty nine turn left heading three six zero.

1225:59 {03:42}

RDO-2

left three sixty, AC five twenty nine.

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1226:03 {03:46}

HOT-1

no he didn't, didn't know whether, the high pack was causing the

problem or.

1227:29 {05:12}

ATLD

AC five twenty nine turn left heading two seven zero.

1227:33 {05:16}

RDO-2

left * two seven zero, AC five twenty nine.

1227:41 {05:24}

HOT-1

let's look at the radar there.

1227:44 {05:27}

HOT-2

just a bunch of scattered stuff *.

1228:37 {06:20}

ATLD

AC five twenty nine, maintain one one thousand.

1228:40 {06:23}

RDO-2

one one thousand, AC five twenty nine.

1228:42 {06:25}

HOT-2

eleven.

1228:43 {06:26}

HOT-1

1228:59 {06:42}

HOT-1

stick it back in auto and see if it'll, work again.

1229:35 {07:18}

HOT-2

one more. (feels) alright.

1230:05 {07:48}

CAM

[sound of three beeps similar to altitude alert signal]

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1231:33 {09:16}

ATLD

AC five twenty nine, turn left heading two five zero.

1231:35 {09:18}

RDO-2

two five zero, AC five twenty nine.

1232:02 {09:45}

ATLD

AC five twenty nine, climb and maintain one two thousand.

1232:06 {09:49}

RDO-2

one two thousand, AC five twenty nine.

1232:08 {09:51}

HOT-2

twelve.

1232:09 {09:52}

HOT-1

tell Robin it just be a couple of minutes it'll smooth out.

1232:18 {10:01}

CAM

[sound of two chimes similar to flight attendant call chime]

1232:19 {10:02}

INT-3

hello.

1232:20 {10:03}

INT-2

hey Robin.

1232:21 {10:04}

INT-3

hi.

1232:22 {10:05}

INT-2

it'll be a just a couple more minutes like this, it'll smooth out.

1232:24 {10:07}

INT-3

uuh, couple more minutes and then I can get up?

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1232:25 {10:08}

INT-2

yes ma'am.

1232:26 {10:09}

INT-3

alright, thank you.

1232:27 {10:10}

INT-2

see ya.

1232:40 {10:23}

HOT-2

four more.

1232:41 {10:24}

HOT-B

[three beeps similar to altitude alert signal]

1234:52 {12:35}

ATLD

AC five twenty nine, climb and maintain one three thousand.

1234:55 {12:38}

RDO-2

one three thousand, AC five twenty nine.

1234:57 {12:40}

HOT-1

thirteen.

1235:01 {12:44}

HOT-1

props ninety.

1235:02 {12:45}

HOT-?

***.

1235:07 {12:50}

HOT-1

[mechanical vibrating sound starts similar to boom microphone vi-

brating against cockpit surface]

1235:15 {12:58}

CAM

[sound similar to propellers decreasing in RPM]

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1235:31 {13:14}

HOT-B

[sound of three beeps similar to altitude alert]

1235:34 {13:17}

CAM-2

four more.

1235:42 {13:25}

CAM-1

** lights out *****.

1235:54 {13:37}

PA-2

ladies and gentlemen, good afternoon, welcome aboard Atlantic

Southeast Airlines flight seventy five twenty nine, service to Gulfport.

we're passing through thirteen thousand feet, captain has turned off

the fasten seat belt sign. you are free to move about the cabin as

you wish. however if you're in your seats, we ask you do so with your

belts fastened loosely around you, just in case we encounter any tur-

bulence enroute. Gulfport on the hour is calling for partly cloudy

skies, temperature of eighty degrees, and winds ten miles an hour

out of the northeast. there's anything we can do to make your flight

more enjoyable, please do not hesitate to call upon us. and thank

you for flying with ASA.

1235:58 {13:41}

ATLD

AC five twenty nine maintain one four thousand. contact At-

lanta center on one three four point niner five.

1236:05 {13:48}

RDO-1

one three four ninety five and fourteen thousand, AC uh, five

twenty nine.

1236:10 {13:53}

ATLD

good day sir, it's thirty four ninety five.

1236:13 {13:56}

RDO-1

thirty four ninety five.

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1236:20 {14:03}

HOT-B

[single beep similar to radio frequency change]

1236:25 {14:08}

RDO-1

center, AC five twenty nine is out of thirteen for fourteen.

1236:28 {14:11}

CTR

AC five twenty nine, Atlanta center roger. I'll have a higher al-

titude for you in just a moment.

1236:32 {14:15}

RDO-1

five twenty nine.

1236:35 {14:18}

CAM-2

I'm back.

1236:36 {14:19}

CAM-1

alright, up to fourteen and, talking to center, expectin' higher in a

while.

1236:42 {14:25}

CAM-2

OK.

1237:10 {14:53}

HOT-B

[sound of three beeps similar altitude alert signal]

1237:14 {14:57}

CAM-2

***.

1237:20 {15:03}

CAM-1

yeah, I guess. I'm not sure what the hell is this?

1237:32 {15:15}

CTR

AC five twenty nine, climb and maintain one five thousand.

1237:34 {15:17}

RDO-2

one five thousand, AC five twenty nine.

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1237:38 {15:21}

CAM-2

*** fifteen ***.

1237:51 {15:34}

CAM-1

**** I can't get to.

1238:10 {15:53}

CAM-1

something underneath the #.

1238:17 {16:00}

CAM-2

it'll drive you nuts.

1238:18 {16:01}

CAM-1

it will drive you nuts.

1238:29 {16:12}

HOT-B

[three beeps similar to altitude alert]

1238:32 {16:15}

CAM-2

four more.

1239:24 {17:07}

CTR

AC five twenty nine, climb and maintain flight level one niner

zero.

1239:27 {17:10}

RDO-2

one niner zero, AC five twenty nine.

1239:30 {17:13}

CAM

[single beep similar to altitude alert]

1241:32 {19:15}

CAM-1

**long* duct tape ** around here *****.

1241:57 {19:40}

CTR

AC five twenty nine, climb and maintain flight level two zero

zero.

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1242:01 {19:44}

RDO-2

two zero zero, AC five twenty nine.

1242:04 {19:47}

CAM-2

twenty.

1242:40 {20:23}

CTR

AC five twenty nine, climb and maintain flight level two four

zero.

1242:44 {20:27}

RDO-2

two four zero, AC five twenty nine.

1242:46 {20:29}

CAM

[sound similar to dialing altitude alerter]

1242:50 {20:33}

CAM-2

twenty four.

1242:51 {20:34}

CAM-1

twenty four.

1243:25 {21:08}

CAM

[sound of several thuds]

1243:26 {21:09}

CAM-1

****.

1243:28 {21:11}

CAM

[three chimes similar master warning] auto-pilot, engine control, oil

[and continues to repeat]

1243:29 {21:12}

CAM-?

1243:32 {21:15}

CAM-2

pack off.

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1243:34 {21:17}

CAM-1

1243:38 {21:21}

CAM-1

we got a left engine out. left power lever. flight idle.

1243:45 {21:28}

CAM

[shaking sound starts and continues for thirty three seconds]

1243:46 {21:29}

CAM-1

left condition lever. left condition lever.

1243:48 {21:31}

CAM-2

yeah.

1243:49 {21:32}

CAM-1

feather.

1243:51 {21:34}

HOT-B

[series of rapid beeps for one second similar to engine fire warning]

1243:54 {21:37}

CAM-1

yeah we're feathered. left condition lever, fuel shut-off.

1243:59 {21:42}

CAM-1

I need some help here.

1244:02 {21:45}

CAM

[mechanical voice messages for engine control and oil cease.

chimes and auto-pilot warning continues]

1244:03 {21:46}

CAM-2

OK.

1244:03 {21:46}

CAM-1

I need some help on this.

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1244:05 {21:48}

CAM-?

(you said it's) feathered?

1244:06 {21:49}

CAM-1

uh,

1244:07 {21:50}

CAM-2

it did feather.

1244:07 {21:50}

CAM-1

it's feathered.

1244:09 {21:52}

CAM-2

OK.

1244:09 {21:52}

CAM

[master warning chimes and voice warning continues]

1244:10 {21:53}

CAM-1

what the hell's going on with this thing.

1244:13 {21:56}

CAM-2

I don't know... got this detector inop.

1244:16 {21:59}

CAM-1

OK ***.

1244:18 {22:01}

CAM-?

OK, let's put our head sets on.

1244:20 {22:03}

CAM-1

I can't hold this thing..

1244:23 {22:06}

CAM-1

help me hold it.

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1244:24 {22:07}

HOT-2

OK.

1244:26 {22:09}

CAM-1

alright comin' on headset.

1244:26 {22:09}

RDO-2

Atlanta center. AC five twenty nine, declaring an emergency.

we've had an engine failure. we're out of fourteen two at this

time.

1244:31 {22:14}

CTR

AC five twenty nine, roger, left turn direct Atlanta.

1244:33 {22:16}

HOT-1

# damn.

1244:34 {22:17}

RDO-2

left turn direct Atlanta, AC five twenty nine.

1244:36 {22:19}

HOT-?

[sound of heavy breathing]

1244:41 {22:24}

HOT-?

** back **.

1244:57 {22:40}

HOT-?

[sound of squeal]

1245:01 {22:44}

CAM

[tone similar to master caution cancel button being activated. all

warnings cease]

1245:03 {22:46}

HOT-1

alright turn your speaker off. oh, we got it. its.

1245:07 {22:50}

HOT-1

I pulled the power back.

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1245:10 {22:53}

CTR

AC five twenty nine, say altitude descending to.

1245:12 {22:55}

RDO-2

we're out of eleven six at this time. AC five twenty nine.

1245:17 {23:00}

HOT-1

alright, it's, it's getting more controllable here ... the engine ..... let's

watch our speed.

1245:32 {23:15}

HOT-1

alright, we're trimmed completely here.

1245:38 {23:21}

HOT-2

I'll tell Robin what's goin' on.

1245:39 {23:22}

HOT-1

yeh.

1245:44 {23:27}

HOT-B

[sound of two chimes similar to cabin call button being activated]

1245:45 {23:28}

INT-3

yes sir.

1245:46 {23:29}

INT-2

OK, we had an engine failure Robin. we declared an emergency,

we're diverting back into Atlanta. go ahead and uh, brief the passen-

gers. this will be an emergency landing back in.

1245:55 {23:38}

INT-3

alright. thank you.

1245:56 {23:39}

HOT-1

tell 'em we want ...

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1245:58 {23:41}

CTR

AC five twenty nine, say altitude leaving.

1246:01 {23:44}

RDO-2

AC five twenty nine's out of ten point three at this time.

1246:03 {23:46}

CTR

AC five twenty nine roger, can you level off or do you need to

keep descending?

1246:09 {23:52}

HOT-1

we ca.. we're gonna need to keep con.. descending. we need a air-

port quick.

1246:13 {23:56}

RDO-2

OK, we uh, we're going to need to keep descending. we need

an airport quick and uh, roll the trucks and everything for us.

1246:20 {24:03}

CTR

AC five twenty nine, West Georgia, the regional airport is at

your .. ten o'clock position and about ten miles.

1246:28 {24:11}

RDO-2

understand ten o'clock and ten miles, AC five twenty nine.

1246:30 {24:13}

CTR

's correct.

1246:36 {24:19}

HOT-1

(* give me) [whispered]

1246:38 {24:21}

HOT-1

let's get out the uh .... engine failure check list, please.

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1246:47 {24:30}

HOT-2

OK, I'll do it manually here.

1246:55 {24:38}

HOT-2

OK, engine failure in flight.

1246:57 {24:40}

CTR

AC five twenty nine, say heading.

1246:59 {24:42}

RDO-2

turnin' to about uh, three ten right now.

1247:01 {24:44}

HOT-2

power level's, flight idle.

1247:03 {24:46}

CTR

AC five twenty nine, roger. you need to be on about a zero

three zero heading for West Georgia Regional, sir.

1247:07 {24:50}

RDO-2

roger, we'll ("probly", or possibly,"try'ta") turn right. we're having

uh, difficulty controlling right now.

1247:11 {24:54}

HOT-2

OK, condition lever's, feather,

1247:13 {24:56}

HOT-1

alright.

1247:14 {24:57}

HOT-2

it did feather... Np's showing zero.

1247:18 {25:01}

HOT-1

'K.

1247:19 {25:02}

HOT-2

OK.

100

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1247:20 {25:03}

CTR

AC five twenty nine, when you can, it's zero four zero.

1247:22 {25:05}

RDO-2

zero four zero, AC five twenty nine.

1247:25 {25:08}

HOT-2

'K, electric, yeah OK it did feather. there's no fire.

1247:27 {25:10}

HOT-1

alright.

1247:28 {25:11}

HOT-2

OK..

1247:32 {25:15}

HOT-2

main auxiliary generators of the failed engine off.

1247:35 {25:18}

HOT-1

'K. I got that.

1247:40 {25:23}

HOT-2

'K, APU .. if available, start. want me to start it?

1247:45 {25:28}

HOT-1

we gotta, bring this down, bring those. put the that off. bring the ice

off...

1247:54 {25:37}

HOT-B

[sound of chime similar to master caution starts and repeats at six

second intervals until the end of the recording]

1247:56 {25:39}

HOT-?

1247:56 {25:39}

CTR

AC five twenty nine uh, say your altitude now sir.

101

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1247:59 {25:42}

RDO-2

out of seven thousand, AC five twenty nine.

1248:00 {25:43}

HOT-B

[sound of three chimes followed by voice message] trim fail. [warning

starts and continues]

1248:04 {25:47}

HOT-1

good start.

1248:04 {25:47}

CTR

AC five twenty nine, I missed that, I'm sorry.

1248:06 {25:49}

RDO-2

we're outta six point nine right now, AC five twenty nine.

1248:09 {25:52}

CTR

AC five twenty nine roger, West Georgia Regional, heading

zero seven zero.

1248:13 {25:56}

RDO-2

zero seven zero, AC five twenty nine.

1248:20 {26:03}

HOT-B

[sound of single beep]

1248:33 {26:16}

HOT-2

OK, it's up and running, Ed.

1248:34 {26:17}

HOT-1

alright, go ahead.

1248:35 {26:18}

CTR

AC five twenty nine, West Georgia Regional is your closest air-

port. the other one's uh, Anniston and that's about thirty miles

to your west, sir.

102

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1248:40 {26:23}

HOT-1

how long, how far West Georgia Reg ... what kind of a runway they

got?

1248:44 {26:27}

RDO-2

what kind of runway's West Georgia Regional got?

1248:54 {26:37}

HOT-1

go ahead and finish the check list.

1248:58 {26:41}

CTR

West Georgia Regional is uh, five say one six and three four

and it's five thousand feet ....

1249:01 {26:44}

HOT-2

OK, APU started. OK, prop sync, off. prop sync's comin' off.

1249:03 {26:46}

HOT-1

OK.

1249:04 {26:47}

HOT-2

fuel pumps failed engine. you want uh, max on this?

1249:07 {26:50}

HOT-1

go ahead, please.

1249:08 {26:51}

HOT-2

OK.

1249:09 {26:52}

CAM

[sound similar to propeller increasing in RPM]

1249:09 {26:52}

CTR

... and it is asphalt sir.

1249:11 {26:54}

HOT-2

hydraulic pump, failed engine? as required. put it to the on position?

103

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1249:15 {26:58}

HOT-1

correct.

1249:17 {27:00}

HOT-2

'K. engine bleed failed engine is closed and the pack is off.

1249:19 {27:02}

HOT-1

'K.

1249:26 {27:09}

HOT-2

'K, cross-bleed open.

1249:29 {27:12}

HOT-1

'K.

1249:32 {27:15}

HOT-2

electrical load, below four hundred amps.

1249:38 {27:21}

HOT-1

it is. put the ice ba ... (well you) don't need to do that just leave that

alone.

1249:45 {27:28}

HOT-1

alright, single engine check list please.

1249:48 {27:31}

CTR

AC five twenty nine, I've lost your transponder. say altitude.

1249:52 {27:35}

RDO-2

we're out of four point five at this time.

1249:54 {27:37}

CTR

AC five twenty nine, I've got you now and the airport's at your,

say say your heading now sir.

104

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1249:59 {27:42}

RDO-2

right now we're heading uh, zero eight zero.

1250:01 {27:44}

CTR

roger, you need about ten degrees left. should be twelve

o'clock and about eight miles.

1250:05 {27:48}

RDO-2

ten left, twelve 'n eight miles and uh, do we got a, ILS to this

runway?

1250:10 {27:53}

CTR

uh, I'll tell you what. let me put you on approach. he works that

airport and he will be able to give you more information. con-

tact Atlanta approach on one two one point zero, sir.

1250:15 {27:58}

HOT-1

we can get in on a visual.

1250:17 {28:00}

RDO-2

one more time on the freq..

1250:20 {28:03}

RDO-1

say again the frequency?

1250:22 {28:05}

CTR

Atlanta approach one two one point zero.

1250:24 {28:07}

RDO-2

twenty one zero, see ya.

1250:26 {28:09}

UNK-?

good luck guys.

1250:27 {28:10}

RDO-2

'preciate it.

105

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1250:28 {28:11}

HOT-B

[single beep similar to radio frequency change]

1250:29 {28:12}

RDO-2

Atlanta approach, AC five twenty nine's with you out of three

point four.

1250:36 {28:19}

HOT-1

engine's exploded. it's just hanging out there.

1250:43 {28:26}

RDO-2

Atlanta approach, AC five twenty nine.

1250:45 {28:28}

ATLA

AC five twenty nine, Atlanta approach.

1250:48 {28:31}

RDO-2

yes sir, we're with you declaring an emergency.

1250:49 {28:32}

ATLA

AC five twenty nine, roger. expect localizer runway three four

approach and uh, could you fly heading one eight zero uh no

sorry, one six zero?

1250:56 {28:39}

RDO-2

yeah we can do that. give me the loc freq ...

1250:59 {28:42}

ATLA

localizer frequency, runway three four localizer frequency is uh,

one one one point seven.

1251:05 {28:48}

HOT-1

we can get in on a visual. just give us vectors.

1251:07 {28:50}

RDO-2

one one one point seven ..... just give us vectors. we'll go the

visual.

106

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1251:17 {29:00}

HOT-1

sing, single, single engine check list, please.

1251:28 {29:11}

HOT-2

where the # is it?

1251:29 {29:12}

ATLA

AC five twenty nine, say altitude leaving.

1251:31 {29:14}

RDO-2

we're out of nineteen hundred at this time.

1251:33 {29:16}

HOT-1

we're below the clouds. tell 'm ...

1251:35 {29:18}

ATLA

you're out of nineteen hundred now?

1251:36 {29:19}

RDO-2

'K we're uh, VFR at this time. give us a vector to the airport.

1251:39 {29:22}

ATLA

AC five twenty nine. turn left uh, fly heading zero four zero.

bear, the uh, airport's at your about ten o'clock and six miles sir.

radar contact lost this time.

1251:47 {29:30}

RDO-2

zero four zero, AC five twenty nine.

1252:07 {29:50}

HOT-M

five hundred.

1252:10 {29:53}

HOT-M

too low gear. [starts and repeats]

107

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1252:11 {29:54}

ATLA

AC five twenty nine, if able, change to my frequency, one one

eight point seven. the airport uh, in the vicinity of your ten

o'clock at twelve o'clock and about four miles or so.

1252:20 {30:03}

HOT-1

help me, help me hold it, help me hold, help me hold it.

1252:26 {30:09}

ATLA

AC five twenty nine, change frequency, one one eight point

seven if able.

1252:32 {30:15}

HOT-B

too low gear [warning stops]

1252:32 {30:15}

HOT-B

[series of rapid beeps similar to aural stall warning]

1252:32 {30:15}

CAM

[vibrating sound similar to aircraft stick shaker starts and continues

for four seconds]

1252:36 {30:19}

CAM

[vibrating sound similar to aircraft stick shaker starts again and con-

tinues to impact]

1252:37 {30:20}

HOT-2

Amy, I love you.

1252:40 {30:23}

HOT-B

landing gear.

1252:41 {30:24}

CAM-?

[sound of grunting]

1252:45 {30:28}

CAM

[sound of impact]

108

INTRA-COCKPIT COMMUNICATION AIR-GROUND COMMUNICATION

TIME & TIME &

SOURCE CONTENT SOURCE CONTENT

1252:46 {30:29}

HOT-B

landing gear

1252:46 {30:29}

CAM

[sound of impact]

1252:46 {30:29}

END of RECORDING

END of TRANSCRIPT

109

110

APPENDIX C

Aircraft/

Blade

Model

ATR42/

14SF-5

EMB120/

14RF-9

FRACTURE SUMMARY *

Fracture

Operator

Location

Initiating

Defect

InterCanadian

Taperbore

0.031" deep pit

0.058"

wide

Nordeste

Taperbore

Band

pits .011 -.015"

deep x 0.160" wide

EMB120/

ASA

14RF-9

Taperbore

Pit (initial size

unknown);

.005"deep x 0.037"

wide

(at

surface)

0.011 wide (subsurface)

12,038

4,748

4,210

N/A

14,664

2,399

Blade

Observations

S/N

—

Coarse

banding

856922

10,000-15,000

Oxidized beach mark;

865093

0.032"deep x

0.160" wide

Oxidized beach mark;

861398

0.0487

deep x

0.0542"

wide

(at

surface)

0.066"

wide [subsurface)

All dimensions after rework

*Information provided by Hamilton Standard for

February 2, 1996, briefing to the Safety Board.

**The Safety Board’s investigation revealed that the total time

was

14,728

operating hours and 5,182

hours since overhaul.

10.

111

APPENDIX D

EMB-120/14RF-9

STRESS RESURVEY *

Summary

Flight Test Results

The

highest

vibratory stress

measured in flight was

±6000 psi and

occurred at the

20-inch and the

26-inch

stations during maneuvers

outside the normal flight envelope.

Vibratory stresses measured on the left propeller were about 2%

higher than those seen in the 1985 survey.

Vibratory stresses measured on the right propeller were about

higher than those seen in the 1985

survey.

The right propeller stresses were about 6% higher than the left.

Some small differences in stressing amongst the blades were

observed. Generally, blade number four was the highest.

Average

relative levels for the other blades were: Blade no.

1- 97%

Blade

no. 2 - 95%, Blade no. 3-

96%, Blade no. 5-

100%. Also, there

was a standard deviation of

±3%

on these relationships.

The frequency of vibratory stressing for

all

the strain gages excep

the shank leading edge was primarily 1P. The shank leading edge

gage was primarily 2P with

also present during high-stress

operation. This is consistent with historical data.

The Vee

gage was

always low in magnitude indicating no

significant torsional response. This was as expected.

The distribution of stress along the blade was seen to be slightly

different than theoretical. The 20-inch station was sometimes

higher than the 26-inch station. Analysis indicates that the

26-inch

station should be the highest. Also, the reduction in stress going

inboard from the peak is less than theoretical, see plot.

The trend of stress versus

aircraft

c.g.

vertical

acceleration

compares favorably with previous data, see plots.

The stresses experienced at the extremes of yaw maneuvers were

considerably

lower than those seen in the previous tests.

*Information provided by Hamilton Standard for a February 2, 1996, briefing to the

Safety

Board.

112

Summary of Ground Test Results

The highest vibratory stresses of the entire test

occurred

during

ground

running

in an adverse wind when the propeller speed

passed through the 2P/1f critical. The peak stress reached

±18,600 psi at the 26-inch station of blade number four. The wind

was from the rear quarter at approximately 28 knots.

The 2P/1f critical speed was found to exist at approximately 935

rpm, which is between the locations observed in the previous

two

tests (1983 and 1985), see plots.

The

vibratory stressing that occurs under these conditions is

almost completely 2P in frequency. Adjacent blades are out of

phase such that the reactions of the four blades cancel within the

hub. This is

consistent

with previous data.

Frequency analysis of the strain gage data showed poorly defined

peaks in the vicinity of the expected natural frequency. This could

be due to mis-tunning or other unknown reasons.

The Vee gage was always low in magnitude indicating no

significant torsional response. This was as expected.

The distribution of stress along the blade was seen to be slightly

different from previous experience, see plot.

The high stressing associated with the 2P/1f critical speed were

only observed during adverse wind operation. Tests in mild

headwinds showed very low levels of stress. Again this is as

expected.

In addition to the high stresses seen during operation in the region

of the 2P/1f critical speed, additional stress peaks were

observed

during transient conditions at speeds other than the critical. These

peaks appear to happen immediately after a power lever

movement, independent of the operating speed. They look exactly

like a rapid excursion through the critical speed, 2P in frequency

with adiacent blades out of

phase,

see

examples.

113

Conclusions & Recommendations

The

results

of this stress

survey

are consistent with past experience.

Small differences from previous stress levels were

observed.

Also,

differences were seen among the blades and between

the left

and

right

nacelles. These differences are not considered significant

The highest

vibratory

stressing occurred during

ground

operation in

adverse winds when the

prope!ler

speed was at

near 935 rpm

(72%

take-off).

This is due to the presence of a critical speed caused by

the coincidence of the

first

flatwise

blade mode and excitation due to

the second harmonic of the rotational

speed.

During the worst of

these conditions the vibratory stresses

were more

than

triple the

highest values measured

flight, including

maneuvers.

Operation

this

speed range must be kept to an absolute minimum to avoid

fatigue damage to the blades.