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Launched Aug 26 1996.


Another example of one accident, two investigation and two very different outcomes.

This report describes Petitioner's ivestigation of a 1991 accident which produced very different findings than those reported by the National Transportation Safety Board; in its report of its investigaiton of the same accident - NTSB/AAR-93/O1/SUM, "Loss of Control, Business Express, Inc., Beechcraft 1900C N811BE Near Block Island Rhode Island, December 28, 1991" PB93-910405. That report should be read first.

From an investigation research perspective, the work on the auidio recording tapes is of particular interest, in that it suggests investigators need to be aware of the potential for hidden data in any diata sources during investigations. Both scenarios are deduced from information produced by the accident, but developed by different investigation methods and techniques. The comparison of the human factors, management and training conclusions should be of special interest to investigation process researchers.

As of this date, the NTSB has not responded to this ALPA Petition. When it does, the response will also be posted with these two reports. Until then, you might wish to ponder the consequences suggested by this set of differing reports.

L Benner

535 Hemdon Parkway P.O.Box 1169, Hemdon VA 22070-9805 (703) 689-2270

Petition for
of Probable Cause


Business Express,
N811BE, SN UB-49
Block Island, Rhode Island
December 28, 1991
Accident No. NYC-92-FA-053

June 25, 1997

Submitted by
James M. Walters
Steven D. Green,
Airline Pilots Association


Table of Contents

(The following were not reproduced for this posting)
Diagrams: Wreckage Documentation and Systems Layouts, #1 -
Photographs: Wreckage Photographs, #1 - #54
Attachments: Attachments A - H

Compact Disc: Inside of back cover

List of Attachments

Attachment A:

AIG - Business Express, Inc., SN UB-49 Beech 1900C, Failure Investigation, Packer Engineering, (metallurgy report)

Attachment B:

Radar Data Study, Data List, Plots, N811BE, Beech 1900C, Block Island, Rhode Island, Associated Data Resource, (wreckage plots and drift reports)

Attachment C:

Letters, (2), from Mr. Joseph G. Dondero, Vice President Maintenance, Pennsylvania Airlines, and spokesman for the Beech 1900 Operators Committee, to Mr. Dave Jacobson, Manager of Airline Support, Beech Aircraft Corp., dated September 3, 1991 and October 16, 1991.
FAA Internal Memorandum, Principal Aviation Inspector (Airworthiness) NE-FSDO-03 to Manager, Aircraft Certification Directorate ACE- 100.
Airworthiness Directive 91-12-02
NPRM Airworthiness Directive 93-CE-4 1-AD, sp. AD 92-06-09.


Attachment D:

The Beech Aircraft Corporation Model 1900 Airliner Engine Truss: A Study in Reliability Analysis and Aviation Safety, Dr Ron Stearman P.E., M. Buschow and K Kane
C.V. Dr. Ron Stearman, P.E.


Attachment E:

Aircraft Damage Detection from Acoustic and Noise Impressed Signals Found by a Cockpit Voice Recorder, Dr. Ron Stearman P.E., G. Schulze, and S. Rohre, Institute of Noise Control Eng.


Attachment F:

Materials Engineering Investigation Conclusions, Chronological Event Graph, Dr. Richard H. McSwain, P.E., McSwain Engineering, Inc. C.V. Dr. Richard McSwain, P.E

Attachment G:

Expert Report, Donald Hamill (engine teardown) C.V. Don Hamill

Attachment H:

Cockpit Voice Recorder Tape Erasure Gap Study and Signal Level Inventory and Study, Glen Schulze, Data Acquisition Sys.
. Glen Schulze

page ii



On December 28, 1991, a Beech Aircraft Corp. 1900C, operated by Business Express Inc. (BEX) crashed while executing a VOR approach to Block Island airport, Rhode Island. The flight was operated under the provisions of 14 Code of Federal Regulations (CFR) Part 91, and under visual flight rules (VER). As a training flight, only company personnel were on board. There were three fatalities, including the instructor pilot and two trainees. The National Transportation Safety Board (NTSB) determined the probable cause of this accident to be:

“...the instructor pilot’s loss of altitude awareness and possible spatial disorientation, which resulted in the loss of control of the airplane at an altitude too low for recovery; and company management’s lack of involvement in and oversight of its Beechcraft 1900 flight training program. Contributing to the accident was the instructor pilot’s exercise of poor judgment in establishing a flight situation and airplane configuration conducive to spatial disorientation that afforded the pilots little or no margin for error.”

The Air Line Pilots Association (ALPA) takes issue with the probable cause adopted by the Board and with numerous statements contained in the Board’s report [1] . In addition, ALPA takes exception to seven of the eleven Findings in the Board’s final report, and believes these findings are erroneous. Those findings, and the errors they contain, in summary, are:

Board finding #2.

“There was no evidence of airframe or powerplant failures prior to impact with the water.”

Errors contained in Board finding #2.

There was (and still is) extensive and irrefutable evidence that, in fact, there was structural failure of several key components of the accident aircraft while maneuvering at approxnuately 2000’ altitude . Those would include the right engine truss tube assembly, which caused a whirl flutter event in the right engine and propeller and subsequent wing and empennage failures. AU occurred prior to impact with the water.

Board finding #3.

“There were no airplane system malfunctions or failures before impact with the water, except when electrical power to the captain-trainee’s attitude indicator was deliberately removed.”


Page 1

Errors contained in Board finding #3.

As stated above, ALPA has documented evidence supporting the scenario of a catastrophic inflight breakup. Obviously, during and after this breakup, there were disruptions and failures of all aircraft systems.

Board finding #5.

“The IP [Instructor Pilot] disabled the captain-trainee’s attitude indicator, and about 6 minutes later he simulated a failure of the right engine by retarding the power lever to the flight idle position, which in effect, introduced multiple emergencies contrary to the provisions of the company’s BE 1900 operating manual.”

Errors contained in Board finding #5.

At no time were any company operations specifications or operating policies and procedures violated by the instructor pilot or trainees. The IF simulated the failure of the right engine by reducing the power to zero thrust only , not to flight idle. The IP had valid reasons for simulating those particular system failures during this training fight.

Board finding #6.

‘The IP used poor judgment by encouraging the captain-trainee to fly with his attitude indicator disabled and uncovered, followed about 6 minutes later by a simulated failure of the right engine under simulated instrument conditions on a dark night.”

Errors contained in Board finding #6.

The IF demonstrated his technical and professional competence during all training maneuvers, as documented on the CVR.

Board finding #7.

The IP failed to recognize in a timely manner that the captain-trainee was spatially disoriented when the captain-trainee asked the IP to take control of the airplane (‘Your aircraft?”); instead the IP attempted to coach the captain-trainee into a recovery from an unusual attitude.

Errors contained in Board finding #7.

There is no evidence that the IP was ever “spatially disoriented” during the flight. There is substantial evidence that he was continually aware of the


Page 2


attitude and altitude of the airplane at all times. Although bank angle may have been excessive during the last few seconds of controlled flight, there is no evidence that the trainee was ever in an “unusual attitude”. The comment “Your airplane?” may in fact be a query as to who should be or actually was flying the aircraft, not a request for the IP to take over control.

Board finding #8.

The attempted recovery from an unusual attitude was not successful, apparently because the IP lost awareness of the airplane’s altitude and rates of descent and may have become spatially disoriented at an altitude too low for recovery.”

Errors in Board finding #8.

The IP never lost awareness of the airplane’s altitude or rate of descent and never became spatially disoriented. Therefore, because no “unusual attitude” was encountered, there was never a recovery attempt. The last few minutes of radar returns, including the very last return, show the aircraft in level flight at 1900 feet. In fact, due to whirl mode flutter and pre-existing structural faults, the right engine departed the aircraft in level flight, striking and removing the right horizontal stabilizer. This caused the instantaneous failure of the outboard portions of both wings, and the total loss of control of the aircraft.

Board finding #9.

“The airplane probably crashed into the ocean in a near-inverted attitude with the outboard section of the left wing striking the water first and with the longitudinal axis at a substantial angle with respect to the surface of the ocean.”

Errors in Board finding #9.

The exact attitude of the aircraft at time of impact with the ocean is unknown. However, a major portion of the right wing, the right engine & nacelle, the right horizontal stabilizer and the outboard portion of the left wing were no longer attached to the airframe upon water impact, having separated at approximately 1900 feet MSL. It is believed, based on documentation done by the Board as well as that done by ALPA, that the remainder of the aircraft impacted the water nose and left wing root first, in a nearly vertical attitude.

Because it has been over five years since this accident, ALPA and others have had the “luxury” of time to investigate, document and analyze all the factual materials and information related to this inflight breakup. This evidence is obvious and extensive. Much of it has been discovered subsequent to the publishing of the final report by the NTSB. However, ALPA is disappointed


Page 3


that the Board investigator assigned to this accident chose to conduct only a very cursory investigation. During the field phase, many very important facts went undiscovered, and evidence that did not fit neatly into the final scenario was simply overlooked.

New evidence that is presented includes:

  • Detailed history of the well known problem of B 1900 engine truss tube assembly cracking and separation,
  • Wreckage documentation of the landing gear system, the hydraulic system, the elevator trim system and the right engine,
  • Structural documentation of the right nacelle, the lower spar cap and the empennage nose cone,
  • CVR spectroanalysis by two independent laboratories regarding several portions of the CVR tape,
  • Operational considerations, including the use and analysis of the altitude alerter system.

ALPA’s supporting documentation for all assertions is contained in the following petition. The major subject areas are:

Wreckage Documentation
Human Performance
Summary Findings

As required by the Board, new evidence is presented and the errors in the Board’s original report are detailed in ALPA’s Petition for Reconsideration of Probable Cause, Beechcraft 1900C, N811BE.
Page 4


Wreckage Documentation


The NTSB summary report of this accident, NTSB/AAR-93/0 1/SUM adopted April 23, 1993, devotes only two pages to the documentation of the wreckage ofN811BE. Detailed studies of those components of the aircraft that were recovered, and are still in storage today, have been completed subsequent to the publishing of the official report [2] . The findings are as follows:

Right Wing

The right wing failed near WS 124, at the outboard side of the nacelle (see Diagram #1 for locations of wing structural failures, see photos #1, #2 and #3). The lower main spar cap failed at three locations, with the structural member of the lower spar cap separating along the lateral bonded surface:

The after portion of this member failed at WS 149 (failure “F4”, diagram #1).
The forward portion failed further out, at about WS 158 (failure “F5”, diagram #1).
The skin covering the spar cap failed even further out, at WS 211 (failure “F6”, diagram
#1, detail of photo #2).

The upper main spar cap extends inboard from the main fracture site approximately 28 inches. This would place the failure at WS 96, inside the nacelle (failure “F7”, diagram #1, see photo #4 and detail in photo #2). The aft spar failed at the WS 120 connector. However, the upper cap showed a fracture along the lateral axis, behind the web, which apparently initiated at a rivet hole. The main spar web is compression buckled outward to the location of the principal lower cap separation, that is, WS 149 (“F4”, diagram #1, see photo #6). The bottom panel is also compression buckled to approximately the same location.

The Packer Engineering report [3] (attachment A) identifies all of these fractures as tensile overload or tensile tearing fractures. In the case of”F4” and “F7”, diagram #1, rubbed areas are visible.

The control surfaces attached to this wing section are the outboard flap and the aileron. The root of the outboard flap corresponds with the fracture line at WS 124, and this component shows considerable compression buckling (see photo #6). Both the flap and the aileron have been shifted laterally outboard, towards the tip of the wing. This is confirmed by examination of the control surface hinges. All show an outboard displacement.

Photo #5 shows the fracture of the forward lamination strip of the lower main spar cap. This
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segment is twisted approximately 20 degrees out of plane with the aft segment of the cap. This twist corresponds to a leading edge lift of the wing.

Visual examination of the wing shows it to be straight, with no obvious evidence of set. There is no damage whatsoever to the leading edge or wingtip (see photos #7 and #8). The navigation light plastic cover on the right wing is intact. The bulb was removed by the NTSB for microscopic filament examination and was found to be stretched but not broken. The wing is straight, and the skin shows no damage. The stall fence is intact, indicating that the wing could not have struck the water with any speed. The right wing is, in summary, remarkably free of damage! (See photo #9).

As documented in the NTSB report, the right wing was found floating on the. surface of the water over 4 miles from the main wreckage area. It is significant to note that the location of the wing when found was opposite the prevailing direction of surface drift (from the Southeast to the Northwest) but corresponding to the general wind direction (280 degrees at 12 knots) present at the time of the accident (see Coast Guard Logs contained in the “Radar Data Study” [4] and graphic layout, attachment B).

A reasonable conclusion, then, based on all of the preceding evidence, would be that the right wing failed in flight and separated from the rest of the airframe prior to the aircraft entering the water.

Furthermore, it is clear that the right wing failed in downward bending, based principally on the buckling of the main spar web and bottom panel skin in the area where the lower spar cap was separated. It is important to note that the spar structures show no indication of fore/aft bending; the forces applied at the time the spars failed appear to have been strictly vertical. Given this vertical load, the fracture at “F7”, diagram #1 is interesting because it is located at the end of a 28 inch long segment of upper spar cap (see detail in photos #2, #4 and #9). This cap section runs directly under some major bulkheads in the nacelle as well as under the firewall assembly and the cowling (see diagram #2). For it to lift out of the top of the wing as the outer panel failed downward, yet remain true to the wing plane, would require that the nacelle and cowling assemblies be absent. Were they in place, this length of unreinforced spar cap would have encountered the nacelle and cowling, thus causing major damage. No such damage is evident.

Consequently, the nacelle structure departed the upper surface of the wing prior to the wing’s failure in downward bending.

The previously referenced Packer Engineering metallurgy report (attachment A) identified several elongated rivet holes in the wing-to-nacelle skin immediately aft of the upper spar cap. These holes are elongated upward and aft. This indicates that the portion of the nacelle structure attached to this piece of skin was lifted up and moved aft relative to the wing. All or most of the nacelle would have separated with it.

Page 6


So it appears that the right engine/nacelle structure was torn upward and aft relative to the wing, and that the right wing subsequently failed in the downward direction.

Left Wing

The left wing failed in at least two locations. The outboard section, which extends from the tip inboard to the leading edge failure at WS 221 (see photos #10 and #11), includes the main spar failure between WS 216 and WS 221, the trailing edge failure at WS 206, and the complete aileron (photo #12). While the wing failed outboard of the inboard end of the aileron, the aileron itself remained intact and separated with the wing section as a complete piece. This section shows only moderate damage. The leading edge skin between WS 221 and WS 269 is compressed upward and torn from the spar cap (photo #13). However, this area of damage does not show evidence of a high-energy water impact; the exposed spar cap is not at all peeled or deformed. The tip is mangled. The three hinges are shifted aileron outboard. The root of the aileron is compression buckled as is the outboard end of the outboard flap (see photos #14 and #15). The inboard end of the aileron shows obvious deflection and deformation in the downward direction (see photo #16).

The left wing inboard section extends from the WS 210 inboard to and including the main landing gear box with the gear still attached (see photo #17). The leading edge has disintegrated back to the main spar line (photo #18). A section of the main spar web is present, extending from about WS 93 outboard to approximately WS 165, and considerable deformation is present. Both spar caps have separated from this web section. Outboard of this web section, extending to the outboard fracture at WS 210, no spar structure at all is evident.

The left wing main spar failed between WS 216 at the lower cap (“F2”, diagram #1) and WS 221 at the upper cap (“Fl”, diagram #1, photo #18). The lower spar cap fracture at WS 216 mates perfectly with the left end of the recovered lower cap section (see “Packer” photo, attachment A). The upper cap is fractured with no mate recovered. The upper cap fracture site is about five inches outboard of the lower cap fracture site, with the adjacent, uncapped web bowed forward.

It is interesting that each end of the surviving 35 foot section of the lower main spar cap mates with a surviving wing structure. In the area of the “F4” and “F5” site in diagram #1, the spar cap was pulled from the bottom of the wing back to the failure line just outboard of the nacelle. At the “F2” site, the lower cap failure corresponds to the failure line of the left wing outer paneL Both wing failures were in downward bending.

Furthermore, it is interesting to note the difference in the condition of the left wing outer panel and the inner panel (see detail, photo #9). The leading edge of the inner wing section is totally destroyed, with a few sections completely missing. The outer panel, however, is damaged but essentially intact. It does not appear to exhibit any high-energy “hydraulic” damage, that is damage incurred by the action of the panel forcibly striking the surface of the water. Indeed, the upward crush of the leading edge between stations 221 and 269 opens up a perfect scoop along the spar line. Yet the spar cap and skin in this area show no signs of peeling, rolling, or other

Page 7


hydraulic damage (see detail of photo #13). So it is highly unlikely that this section could have entered the water at a high speed, yet highly probable that the inner wing section did.

At Fl of diagram #1, the left wing failed in down bending. The aileron stayed with the outer panel and is complete, which might not be expected in an aft bending separation. There is considerable compression of the trailing inboard tip of the aileron, as well as the leading inboard corner of the wing panel (details of photo #10 and photo #11). These may have resulted from collision during or after separation. Most interesting, however, is the section of skin overlapping the lower spar cap flange just inboard of “F2”, diagram #1. The rivet holes are dimpled but not elongated, showing tension pullout. The skin section is not rolled back, however, and although wrinkled, still lies along the plane of the spar cap. There has been no fore or aft movement during the separation between the lower spar cap and the wing skin in this area, only a very high energy tension “pullout” failure.

In contrast, the inner panel is severely damaged (see photo #9 and photo #18 detail). The leading edge is gone. The upper spar cap is gone. The lower spar cap is part of the 35 foot section recovered intact but separated from the web. A section of spar web remains but is bent fore and. aft. The nacelle is gone. This wing section was probably still associated with the airframe after the major breakup and later, upon impact with the water. The section of skin flange overlapping the lower spar cap and adjacent to the outer panel skin is visible. Here, however, the flange is rolled up and back, and the rivet holes show tearing and elongation, all of which are indicative of “hydraulic” or water impact damage.

In summary, then, the outboard section of the left wing failed in downward bending prior to the airframe’s impact with the water. The inboard section of the wing remained attached to the fuselage throughout the breakup and subsequent impact with the water.

The empennage

The empennage separated from the fuselage at the junction of the vertical fin and the fuselage. The dorsal was not attached; skin damage was present up to canted SS 11.5, where the dorsal would have attached (see photos #19 and #20).

The right horizontal stabilizer was detached from the aircraft, and was never recovered (see diagram #3). Approximately 9.5 inches of the forward spar structure, however, remains attached to the vertical stabilizer. This structure is bent aft nearly 90 degrees against the vertical fin (“F13 diagram #3, photo #21). Approximately 22 inches of the aft spar structure remains attached, and is bent aft 10 to 15 degrees from the normal sweep of the horizontal stabilizer (“F14”, diagram #3, photo #21).

This bending of the remaining spar stubs indicates an aft bending failure of the right horizontal stabilizer. There is no evidence of any vertical axis deformation of these stubs, and the failure appears to have been completely longitudinal.

The outboard portion of the left stabilizer was recovered, extending from approximately HSS 40

Page 8


to the tip. The inboard fracture site area shows compression buckling at the trailing edge (“Fl5”, diagram #3). The two hinges present show severe displacement elevator outboard (photo #22).

The left horizontal stabilizer includes three significant features. The first is that it was physically sawed, by a member of the wreckage recovery team, at approximately HSS 55 (“C 11”, diagram #2). Secondly, the leading edge is characterized by two distinct impact marks. A large circumference depression is centered just inboard of the station 35 rib. This impact caused the leading edge to roll downward and left. A second depression is located at HSS 47, which is a knife-edge indentation approximately six inches long and one inch in width. It is rounded at the bottom, creating a very straight trough. Associated with this trough are several compression buckles in the skin on both the upper and lower surfaces of the stabilizer. Beginning at HSS 20 and extending outboard to this trough, the stabilizer forward spar is bent back approximately eight to nine degrees beyond the nominal (see photos #23, #24, #25, #26 and #27).

The saw line comprises approximately 9 inches of the trough mark. It is now impossible to determine to what extent the stabilizer was damaged prior to the cut. However, the leading edge skin between the strike mark and the severance line is pulled away from the rivets. Several rivet holes are torn in the aft direction. At the adjacent leading edge site of the outboard section, the skin is in place with no separation. Further, the photographs taken at Quonset immediately after recovery show the outer panel of the stabilizer bent substantially downward, not upward as stated in the NTSB accident report (see photos #28 and #29).

Additionally, photo #30 shows a segment of the surface deicing bleed air line mounted in the leading edge of the left horizontal stabilizer. The bleed air line is cut through at the span wise location coincident with the cut in the left horizontal stabilizer. Three additional, smaller cuts are also on this bleed air line.

The vertical stabilizer forward spar is twisted 30 degrees clockwise looking from the top down. Virtually all of this twist occurs prior to SS 11.5 (see photos #19 and #31). This damage appears to be associated with an impact on the right side of the spar just below the SS 11.5.

The forward spar appears to be generally straight above SS 11.5, but the vertical stabilizer aft spar is considerably deformed along its entire length. While the left side of this spar structure is virtually straight, the right cap has a wave set along its length (photos #20 and #32). The web areas that are visible show this wave to be the result of a severe clockwise twist looking from the top down.

The empennage “T-tail” assembly nose cone has a lightweight, aerodynamic fiberglass shell which protrudes forward at the very top of the tail (diagram #2). It is located immediately forward of and adjacent to both horizontal stabilizers, and is physically secured to the top of the vertical stabilizer. This shell was intact, attached in its normal position, and with minimal damage at the time of recovery and subsequent retrieval from the ocean (see photos #23, #31, and #32). Had the previously documented damage to the horizontal and vertical stabilizers occurred during impact with the water, particularly with the aircraft impact attitude assumed by the NTSB, this highly frangible nose cone would have been destroyed. The nose cone’s total lack of damage indicates that all other “T-tail” components received their damage in flight,

Page 9


prior to, and not because of, water impact.

The complete severance of the right horizontal stabilizer represents a fairly high-energy event occurring at the trailing structure. The energy stored by the airframe during water impact would have been expended by the time the tail arrived at the surface, indicating that the damage to the stabilizer occurred prior to the aircraft striking the surface of the water.

In order to produce the observed wave set across the entire length of the vertical stabilizer spar, the two forces involved must be applied at opposite ends of the spar length. The natural candidates for these forces would be the force (impact) that separated the right stabilizer (top of the spar), and the airframe resistance to yaw (at the bottom). Because the forward stabilizer spar stub fractured and is no longer rigidly attached to the aircraft structure, the twist that is seen in the aft spar would not be evident in the forward spar. The aft stubs are still attached to the vertical stabilizer, and would have to have absorbed the entire twisting force, thus “setting” the aft spar in the position observed.

While loss of the right stabilizer was clean and complete, the loss of the left stabilizer must be considered in terms of its failure as a lifting structure. Had it remained intact, it is possible that it may have supported the airframe nose aerodynamically (SA Brasilia at Gadsden). However, if the damage to the leading edge disrupted the airflow over the surface of the airfoil, as the existing damage would have, then the stabilizer’s functionality as an airfoil is reduced to some degree.

The aerodynamic effects of a distortion of the vertical stabilizer are unpredictable, but easily disastrous. In conjunction with the loss of the right stabilizer and damage to the left stabilizer, the airplane at that point was uncontrollable.

Given all of this structural evidence in the empennage, therefore, the following must be concluded; that the right stabilizer was struck by something with enough energy to completely separate it from the tail. Considering the wave set that was imparted to the vertical stabilizer, this impact had to have occurred prior to the empennage separating from the rest of the airframe. Had it occurred after empennage separation, it would simply have implied an angular acceleration to the entire tail, and just spun it around.

Remembering, then, that both wings failed in downward bending in flight, it is most logical to conclude that the right horizontal stabilizer was forcibly struck and subsequently separated from the aircraft in flight. This caused a dramatic reduction in necessary negative lift at the tail of the aircraft, and therefore a severe nose down pitch tendency. Additionally, the partial loss of airflow over the surface of the left stabilizer, in conjunction with the distortion imparted to the vertical stabilizer and the severe nose down trim imparted to the elevator trim system (see documentation and discussion under “Elevator Trim System” page 13 of this report), rendered the aircraft uncontrollable. The extreme nose down pitch tendency caused an instantaneous catastrophic failure of the outboard wing panels in downward bending.

Page 10


Lower Main Spar Cap

A major portion of the lower main spar cap was recovered. At time of recovery from the ocean, it was one long continuous piece. For ease of storage, however, this piece was cut in two (see detail in photo #9). The right hand piece is approximately 19.5 feet in length; the left hand piece is approximately 16 feet in length. The right end begins at the “F3” site, diagram #1. The left end of the lower cap section begins at the “F6” site, diagram #1. The fracture at the “Fl” site and the fracture at the “F6” site appear to be brittle type fractures with little plastic deformation visible around the fracture site. A third fracture is located approximately 9 feet from the left end (“F9”, diagram #1, photos #33, #34 and #3S). This fracture occurred in the downward direction, as in wing down bending. This fracture would have been located close to WS 108, approximately in the center of the left nacelle structure. Notice the darker area on the cap in photo # 33 — this indicates that portion of the spar cap directly under the nacelle. This fracture shows considerable deformation and lipping on either side of the fracture line. Indeed, the spar cap at this location is held together only by the intact segment of the skin.

Also of significance is a distinct scrape mark on the lower surface of the right lower spar cap, at approximately WS 12S (photos #36 and #37).

Right Main Landing Gear

The right main landing gear wheel well and box support structure separated from the rest of the aircraft (see photo #38). The right main gear upper strut support assembly freely rotates around its pivot point, making it impossible to determine the gear upper strut position at time of water impact.

The right main landing gear oleo strut is fractured just above the wheel truck (detail of photo #38, photos #39 and #40). The oleo is partially collapsed above the fracture side, on the forward side of the strut. A triangular scrape mark indicates a sliding impact. The fracture has occurred at the point at which the internal structure of the strut offers the least resistance. This fracture was the direct result of the lower main spar cap separating from the aircraft structure in a severe and rapid downward direction. The relative locations of the right main landing gear oleo strut, while in the “up and locked”, retracted position, and the adjacent position of the lower spar cap, are shown in diagram #2. The distinct witness mark on the lower spar cap (photos #36 and #37) is at the point at which it severed the right main landing gear strut. Note that the mark has a width corresponding exactly with the width of the strut. Note also that photo #36 which shows the unpainted area of the spar cap that passes through the nacelle/wheel well area. The scrape mark and deflection are exactly centered in that area.

The right main landing gear wheels, tires, brake assembly and lower oleo strut separated as a complete unit (see detail of photo #9 and photo #42).

In photo #43, the right main gear assembly has been placed in its relative position to the outboard panel of the right wing. The spar cap area (red in photo) lines up exactly with the fracture in the landing gear oleo strut. See also diagram #2.

Page 11


Therefore, the right main landing gear was in the up and hydraulically “locked” position (see diagram #3) at the time the lower spar cap, right nacelle and engine, and a major portion of the right wing separated violently from the aircraft.

Nose Landing Gear

The nose landing gear drag brace is fractured (photos #44 and #4S). Lips on both ends of the fracture face and uniform bending of the entire segment indicate a column buckling fracture. The degree of deformation on either side of the fracture line is considerable (photo #46).

The forces required for this failure mode are exclusively from forward to aft, causing compression overload failures of the drag brace (photo #47) and the mechanical down lock (see diagram #4). Therefore, at the time of impact with the water, the nose landing gear was down and mechanically locked. This was also the conclusion of the NTSB 10 of the report.

Of obvious significance, then, is the fact that two separate and distinct events occur relative to the position of the landing gear; the main landing gear was up and locked (hydraulically) during the inflight breakup (first event), yet the nose landing gear was down and locked at the time of impact with the water (second event).

The main and nose landing gears are held in the up (locked) position with normal hydraulic pressure through a motor/pump system physically located in the left wing root, immediately forward of the main spar. [5] During an in-flight breakup of the type encountered here, hydraulic pressure would be available to initially keep the gear retracted. However, as electrical power to the 28v motor/pump became unavailable due to numerous electrical overloads, shorts and failures, the only source of normal hydraulic pressure would be lost. Additionally, ruptures and failures of numerous hydraulic lines located in the wheel well and wing spar areas would cause the rapid loss of remaining pressure. That total loss of normal hydraulic pressure would then allow all gear to free fall to the down and mechanically locked position.

The right main gear, however, would by this time have been cut in half, and the response of the short stub of shock strut still attached to the airframe would be unpredictable. The left main gear position would be influenced by the amount and type of deformation caused in the wheel well area by the separation of the outboard portion of the left wing, and portions of the lower spar cap. Indeed, there are indications that the left main gear oleo strut may have been bisected by the rapid “pullout” of the lower spar cap, exactly as happened to the right main gear (see photo #17, and compare with photo #43). Therefore, whether the left main gear would fall completely out of the wheel well, and what position it may have had at impact, is unknown.

However, at some point in the uncontrolled descent and prior to impact with the water, we would expect the nose gear to free fall out of its well to a “drag leg overcenter” or down and locked position. This is exactly what happened, and has been thoroughly documented.

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Elevator Trim

Elevator trim tabs are installed on both elevators. The tabs can be manually controlled by the pilot through drum-cable systems using jackscrew actuators. Moving the cockpit elevator tab control wheel forward results in tab deflection up causing elevator movement down, and aircraft nose down. The amount of wheel and tab movement is shown by reference to a geared position indicator located immediately adjacent to the elevator trim wheel on the cockpit pedestal (see both pages of diagram #5).

The elevator trim tab control cables travel under the left side of the fuselage floor, through pulleys in the fuselage and vertical stabilizer to the tabs located in each elevator (see diagram #5). These cables are fitted with turnbuckles in various locations. One such location is on top of and to either side of the vertical stabilizer, accessible through inspection covers on the top surfaces of the horizontal stabilizers (see detail, diagram #2). These cables and their respective turnbuckles have an operating range (movement along the normal left and right axis of the aircraft, and along the fore/aft axis of the cable) of approximately 8 inches.

Photo #48 shows the pedestal area. Notice the white index mark visible immediately to the right of the trim wheel. Photo #49 is taken from directly above, and is only to document the position of the “nose down” reference beside the trim wheel. The position indicator is geared directly to the elevator trim control wheel. As is evident in this photo, the indicator has been forcibly driven forward, or in the “nose down” trim direction to a point well beyond the normal operating range of the mechanism, but it’s integral drive gearing with the trim control wheel is still intact.

Photos #50, #51 and #52 document the elevator trim control cables as found on the accident aircraft. The cable shown in photos #50 and #51 is the “crossover” cable and turnbuckle, located out in the stabilizer. But note particularly the position of the “nose down” or ‘forward” turnbuckle (see diagram #5, photo #52). Because of pulling action exerted on the right side of the crossover” cable, the “nose down” cable has been pulled in the opposite direction, from it’s normal location inside the right horizontal stabilizer to a position inside the vertical stabilizer, approximately 24 inches away! This would correspond to an elevator trim position far forward of full nose down, which was the post-accident position of the indicator on the cockpit pedestal (photo #49). The cable has had its travel stopped only by the turnbuckle, which is totally inflexible, trying to move around the pulley located in the vertical stabilizer! Once the “nose down” cable had been completely fouled in the pulley, the “crossover” cable was prevented from being pulled any further, and failed outboard of the vertical stabilizer in tensile overload.

As stated earlier, ALPA believes that the damage observed to the nose and left wing root area of the aircraft was caused by initial water impact. Because of normal crash “kinematics”, then, impact damage would occur to the cockpit area prior to the rest of the aircraft, and certainly before the tail, which would have entered the water at a later time than the nose. Therefore, it is evident that the cockpit trim control wheel and indicator have been “back driven” by the cable in the tail. Obviously, the force that would be required to cause movement of the cable along the entire length of the aircraft would have to have originated at the tail . A logical conclusion ,

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then, is that the event causing the “back driving” of the trim system occurred at altitude, prior to the inflight breakup of the aircraft, and long before water impact.

Therefore, the sequence of events would have been:
  • An event at the tail of the aircraft causing extremely rapid movement of the elevator
    pitch trim system towards the “nose down” position,
  • Resultant movement of the elevator trim control wheel and trim indicator in the
    cockpit towards the “nose down” position, only stopping well beyond the full nose down position,
  • At some later time, cockpit impact with the water.

Right Engine Truss

A significant portion of the right engine mounting truss arrived at Pratt & Whitney with the right engine gas generator. This assembly consisted of the hoop structure with the right side V mount still attached (photo #S3). Examination of this structure revealed that of the eight truss tubes attached to the hoop, five had failed at or very near the hoop junctions. A sixth had failed at the hoop junction as well as near the mounting boss. The remaining two, both forward ends of the right side V structure, had not failed.

The aft mounting boss of the right side V structure was present. The truss tube junctions at this boss were intact. The mounting boss itself was closely examined, and the bolt hole itself shows no obvious signs of distress, although there is a scratch visible on the aft side of the boss extending circumferentially about 180 degrees around the bolt hole (photo #54).

Additionally, the V structure is bowed noticeably in the vertical plane. This bow brings the aft mounting boss at the tail of the V inboard relative to the hoop structure (photo #53).

The engine truss structures of the B 1900 have a long history of cracking, often to the point of separation. In 1989 alone, Business Express filed over 70 Service Difficulty Reports with the FAA concerning truss cracks. Other operators have experienced similar difficulties. In September and October of 1991, the “Beech 1900 Operators Committee” was so concerned with this problem that they twice requested action on the part of Beechcraft to remedy this “very serious situation.. and known failure condition” (see attachment C).

But Beechcraft was already aware of the seriousness of the problem, as evidenced in their own summary of the situation. “Numerous instances of fatigue cracks in the Model 1900/1900C engine truss including several instances of complete truss tube separation indicates the potential for degradation of the truss’ load carrying capability.” [6] Beechcraft issued Service Bulletin (SB) 2196 on September 1, 1987, proposing inspections and repair or replacement of cracked truss tubes. Two and a half years later, a mandatory service bulletin (2255) was issued in an attempt to remedy ongoing structural engine mount failures (see chronological sequence of regulatory events, attachment F).

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A detailed and comprehensive reliability study [7] of this problem, completed in June, 1995, is included in its entirety as attachment D.

This report documents 10 years of extensive in service difficulties with all models of B 1900 engine support trusses. Most of the problems have to do with cracking truss tubes, including those resulting in complete separation of the tubes. As early as 1989 the FAA was been aware of the problem, as evidenced by internal memorandums (see attachment D, page 16). In that document, the FAA representative states that because of the engine truss tube structural problems, “the airworthiness of the aircraft will remain in question. This inspector believes this is a very serious safety problem” [8] . Airworthiness Directive (AD) 9 1-12-02 requiring initial and repetitive inspections of the engine trusses and installation of reinforcing doublers was issued May 1S, 1991. A second AD was proposed in November of 1991, and became law three months after the crash of N811BE (both Airworthiness Directives are included in attachment C). Since that time there have been at least three additional AD amendments or revisions and four mandatory Service Bulletin revisions, all pertaining to engine truss tube cracking (see table, attachment F).

237 events of truss tube cracking, separations and failures are documented between 1985 and 1997’ [9] which generated Service Difficulty Reports (DR) (see attachment D, page 17). To date six different truss types have been designed in an attempt to correct the ongoing problem. They vary in tube size, design and thickness, but the increasing mass of the support system has actually decreased the reliability and in-service life expectancy of the truss (see attachment D, pages 18-38).

It is interesting to note that all of the six truss tube failures on the accident aircraft right engine truss occurred at locations that are cited by Beech in four Service Bulletins and two Airworthiness Directives as areas requiring increased inspection.

Please see Compact Disc enclosed (inside backcover) for additional engine truss information. "Whirl Mode Flutter"

The following is excerpted from the previously referenced study by Dr. R.. Stearman, attachment D, pages 9- 11.

“Whirl flutter is the onset of unstable and destructive oscillations usually involving a lifting surface and propeller disk in an airstream. Whirl flutter is a precession-type instability that is the result of the coupling of gyroscopic and aerodynamic forces acting on the propeller. These forces can cause the pitch and yaw degrees of freedom to couple, yielding a whirl mode. Gyroscopic forces can couple the pitch and yaw modes of vibration exhibited in a rotating propeller mounted on a flexible support or structure. This coupling results in one of the following two modes:

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  • forward whirl mode, where the direction of the cyclic precession is the same as the propeller rotation, and
  • backward whirl mode, where the precession is in the opposite direction of the propeller rotation.”.

    The report continues with a detailed discussion of the whirl theory, but concludes that section with “Thus whirl flutter occurs when the aerodynamic forces provide the coupling necessary to induce an unstable whirl motion. Gyroscopic forces tend to destabilize the backward mode further. This phenomenon explains why whirl flutter invariably occurs in the backward mode.

    The report concludes (see attachment D, pages 49-52):
    “Several findings in this investigation suggest that the occurrence of a highly divergent whirl flutter instability led to the destruction of Beech 1900 UIB-49 during the pilot training mission. Some of the more significant findings leading to this conclusion are presented in the following listing:”

    • ‘The statistical reliability study carried out in this investigation indicates a serious fatigue cracking phenomenon within the engine mount truss elements that is getting worse as attempts are made to improve the design. Past experience has shown that damaged engine mount trusses are a known contributor leading to propeller whirl flutter instability of turboprop aircraft.”

    • ‘The engine truss to firewall mounting bolts on UIB-49 could have encountered a foreseeable and significant pre-stress at the time of the truss installation or replacement due to the manufacturing quality control problem illustrated in Figure 3-14. The extent of this bolt and engine truss pre-stress build up will depend upon how the out of tolerance dimensions add up, but will generally, for a given misalignment, be the greatest for the stiffer or newer truss designs. This pre-stress condition will accelerate both fatigue and overload failures. Pre-stressing due to misalignment is not an uncommon occurrence, according to Beech records and a third party informant. This informant suggested that Conquest Airlines of Austin, Texas found it necessary to force fit every replacement engine truss that was installed on the Model 1900C aircraft”.

    • “Further inspection of the right engine truss revealed that it had damage and tube separation in areas where service difficulty reports indicate fatigue cracking regularly occurred...”

    Additionally, both MSC/Nastran 6810 [10] and Brock’s [11] flutter analyses were conducted for this aircraft, engine and propeller configuration. A detailed description of the two studies is included in the report, “Aircraft Damage Detection from Acoustic and Noise Impressed Signals Found by



    a Cockpit Voice Recorder” [12] by Stearman et al, attachment E.

    As documented, when truss cracking occurs in two or more tubes, the whirl flutter speed lies well within the aircraft flight envelope. At 1550 RPM (flight idle), the whirl flutter speed with two broken tubes is precisely in the airspeed range (180k - 190k) in which the accident aircraft was operating.

    An independent study of the engine mounting systems including truss tubes, attaching bolts, nacelle bulkheads, etc. of N811BE was completed subsequent to the accident. [13] This summary document is included in its entirety as attachment F. This study describes the failure mechanism and separation sequence of the right engine mounting truss. As it concludes:

    ‘The right engine mount failed in flight, allowing the engine and engine mount structure to separate from the wing, followed by an in-flight right wing separation. This opinion is based on visual examination of the right wing mount structure and the right wing condition.”

    ‘The failures in the right engine mount truss structure are consistent with propeller whirl mode type loading and the significant Beechcraft 1900 engine mount service failure history has given ample warning of potential engine mount failure and separation as in this accident.”

    Please see enclosed Compact Disc inside back cover for description and demonstration of “whirl mode flutter”


    The official NTSB accident report states that “Both engines displayed scoring from internal rotating parts indicating that the engines were developing power when the airplane struck the water. The scoring was not extensive, and no estimate could be made about the amount of power the engines were developing. However, since one of the last comments on the CVR was ‘power to idle’, followed by sounds of the landing gear warning horn, it appears that the power lever on the left engine was reduced to match the power from the right engine, which was at a flight idle power setting to simulate its failure.” These statements are in error.

    In fact, to simulate the engine failure, the right engine would not have been at flight idle, but at a reduced power setting to simulate zero thrust . Therefore, a number of situations could cause the gear warning horn to sound, including retarding either the left or the right power lever with the gear not down and locked. It is ALPA’s belief, based on discussions later in this report, that the cause of the gear warning horn was, in fact, due to the retarding of the right throttle from the “zero thrust” position to the flight idle position. Furthermore, the actual condition of the

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    engines as found during a very detailed examination done in June of 1995 [14] (see the Hamill report, attachment G) indicates the following:

    * “The left hand engine was developing power at impact as evidenced by the wear
    marks on the rotational parts. This wear and especially heat discoloration on the
    impeller and shroud are indicative of power above low idle even with contact with
    water which has less impact force than solid matter such as the ground or trees.”

    * “The right hand engine was not developing any power at all (emphasis added). The condition of the rotating parts, especially the impeller and shroud, show no signs of wear that can be attributed to the impact force. The impeller and shroud have such a small clearance between them that even low idle power and impact with water would show more than is seen in this instance. I believe that the rotation of the right hand engine was somewhere below low idle at the time of impact. This condition would exist if fuel were not being delivered to the engine for a short period of time prior to impact.”
    The summary section of the Pratt & Whitney engine teardown report [15] states that;

    “The left hand engine and right hand engine gas generator section displayed rotational signatures to the engine internal components characteristic of the engines developing power at impact. The minimal impact deformation of the engine cases limit the severity of the rotational signatures and precludes definitive assessment of the power levels and impact.”

    Additional review of the above referenced report shows that the left engine centrifugal impeller and impeller shroud were “circuniferentially rubbed, with frictional heat discoloration and material transfer, due to axial contact with each other”. The right engine impeller and shroud were only “lightly circumferentially rubbed”, with no heat discoloration or material transfer at all.

    The damage patterns on the gas generator sections show only a positive indication of rotation, whereas the difference in impeller damage between the two engines indicates a difference in power. It is ALPA’s conclusion, then, that the evidence indicates that the left engine was developing power at time of impact, whereas the right engine was not developing power, but only rotating.

    The NTSB goes on to say, ‘The right engine was ... recovered, but during the transfer from the water to the salvage barge, the forward part of the right engine including propeller hub, reduction gears and exhaust casing separated and sank back into the ocean. These latter components were not recovered.” Why the NTSB investigator chose not to recover those portions of the right engine, when their exact location was known, while the recovery team and equipment were m place and in relatively shallow water, is not known.

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    CVR Analysis

    The NTSB provided a transcript of the CVR tape to the investigative team, but at no time was a spectroanalysis study of the recording done by the Safety Board. The official report, page 20, states:
      ‘The Safety Board cannot conclusively exclude the possibility of an event that caused premature termination of CVR operation before the plane struck the water because the CVR events could not be precisely coordinated with the position and altitude as recorded by ATC radar, nor was there any FDR information available to establish the airplane’s actual performance during it’s final descent... .The Board believes that any event that would have caused termination of the CVR must have been sudden and probably catastrophic, which leads to the conclusion that the event was a high speed collision with the surface of the ocean,"
    ALPA agrees with the Board that the event causing the CVR to terminate must have been sudden and catastrophic. But events other than water impact could easily cause cessation of recording.

    Immediately prior to termination of the CVR, as documented by the official transcript and all subsequent CVR studies, there were no recorded acoustic sounds of impact, inflight breakup or vocalization of the pilots indicating awareness of impending disaster. As we will discuss shortly however, there were at least two non-acoustic structural “events” recorded on the CVR, that were detected only by detailed spectroanalysis..

    G Limiting Switch

    The B&D recorder installed on the accident aircraft had a factory installed, 5g limiting switch, which automatically terminates recording with the sensing of a high “g” load upon the switch. This switch prevents inadvertent continual recording by the CVR after an incident or accident, which could erase important data. Any acceleration to the switch of 5g of more, whether imparted to the airframe during the accident sequence or directly to the empennage area (where the CVR is housed) would cause termination of the recording.

    CVR ”Events”

    Several spectroanalysis studies have been conducted on the accident CVR tape subsequent to the Board’s own investigation. The first, the Stearman report (attachment D), labels two very serious events that are detected at the end of the tape, ‘spike 1’ and ‘spike 2’. To quote , “the maximum peaks of the events are approximately 0.263 seconds apart ... The first major spike is preceded by a highly divergent signal of the type present during whirl flutter” (emphasis added). The first spike is immediately preceded by a strong acoustical signature, indicative of a very violent, “explosive” event. The record head shutoff transient occurs at the first spike, while the erase head shutoff transient is evident during the second spike.

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    A power spectrum analysis of the first spike indicates a frequency of approximately 36 Hz. In tests done by Peter Zwillenberg for Beech Aircraft Corporation in Wichita, Kansas [16] , the wing torsion asymmetric vibration mode for the engine in vertical and lateral translation is 37.6 Hz.

    ALPA believes the similarity of the frequency on the CVR and the asymmetric vibration test data to be very significant. If the engine were to “translate”, or tear itself out of the wing, we would expect to see frequencies in the range of 36-37 Hz.

    That same report details the extensive time-series and auto power spectrum analysis that was also conducted. The extensive testing procedures are described in attachment D. To quote only the summary sections,
    .“ . .we suspect that the drastically reduced acoustic power signal associated with the four-blade passage frequency of 104 Hz to 112 Hz (1550 to 1700 rpm) at the end of the tape indicates the loss of about one-half of the source of the signal sound power level - that is, the probable loss of one engine. Regardless of the resolution setting of the analyzer program, (4 or 8 Hz), the results indicated a loss of one-half of the sound power signal. The impact of the engine into the tail would most likely impose an acceleration on the fuselage tail cone structure exceeding the 5 G (five times the force of gravity) threshold required by the CVR for automatic shutoff If the CVR was automatically shut off by an impact with the water, it is likely that there would be noticeable sound of impact (based on the accepted impact scenario) and that both engines would have continued to produce power with the accompanying typical sound power level until impact.”


    “The CVR tape provided some additional strong evidence that a catastrophic event occurred within three to four seconds after the right engine was placed into a flight idle condition. During this operation, the thrust on the right engine would drop to zero and could probably set up a negative thrust or drag configuration on the right propeller disk plane. In [a report for NASA,] ’ [17] Reed demonstrates the stabilizing influence that positive thrust has on the whirl flutter phenomenon. Conversely, a system which might be stable during cruise of normal operating thrust conditions could become unstable when the thrust drops to zero or becomes a negative thrust or drag condition. It is interesting to note that between three to four seconds after a right engine flight idle condition was induced, [the highly divergent acoustical signal is] seen to occur on the CVR tape. This would be characteristic of the aerodynamic lag time for a zero or negative thrusting condition to be induced on the right engine. . . Finally, some of the most convincing evidence on the CVR tape is the loss of 50 % of the engine-propeller sound power level after the two catastrophic events...This clearly indicates that half of the engine-propeller noise source is missing, implying that [an] engine has departed the aircraft. One final factor that is evident from the CVR tape is the fact that the

    page 20

    catastrophic event was so sudden that no one in the cockpit had time to vocally respond. No expletives or other critical comments were noted on the tape. This further suggests a flutter event which is usually an explosive type of phenomenon” (emphasis added).

    CVR “Flutter

    A second, very detailed spectroanalysis study was conducted by Glen Schuize [18] early in 1996 (attachment H). This study isolated and identified a distinct and unique ‘flutter” during the fifth and final beep of the landing gear warning horn. To quote,

    “Figure 30 displays the time-series of a tape track recorded with acoustical voice signals superimposed with the 5 landing gear warning horn signal beeps found right at the end of tape. Amplitude erosion can be seen of the 5th and last horn beep that is not found on the preceding beeps. An attentive listening to these last 5 beeps also revealed a brief but definite audible flutter detected only on the 5th beep.”

    “Figure 31 was obtained at the 5th beep of this horn signal 19 minutes earlier in the tape. No amplitude erosion or audible flutter was found indicating the horn signal source did not suffer fatigue with this length of operational time.”

    In Stearman’s second report, attachment E, he summarizes:

    “Just seconds before the EOT (end of tape), a significant FM modulation was detected both audibly and with the aid of a spectrum analyzer. The modulation was at the propeller fundamental rotation frequency and was due to the dynamic mass unbalance generated by a rotating propeller as it tore loose from its mounting system”.

    Please see the Contact Disk enclosed inside back cover for digital CVR recordings of the “flutter”

    CVR “Silent” Tracks

    One of the four CVR tracks on this aircraft was unused, or “silent”, with no microphone attached.
    This track was found to be rich in non-speech sounds. This process is called electroacoustic transduction, and can include “triboelectric”, "magneto-electric” and “piezoelectric” effects. As Stearman states (attachment E):

    “Close inspection of the amplified (30 dB) times series from the CVR silent track also revealed a very periodic set of transient components occurring at a frequency of 0.86 Hz. Furthermore, this frequency correlated with an independent structural dynamic and flutter analysis of the engine mount damage, which is evident from Figure 3. This transient frequency was found throughout the 32 minutes of this CVR tape indicating the

    Page 21


    condition that generated this phenomenon was an ongoing long-term condition rather than a rapid onset event. These transient components are thought to be typical of impact signatures that would occur from a broken tube end impacting on the tube joint where the fracture occurred. These transients became more frequent about 15 seconds before the airframe disintegrated and then diminished to nearly zero at the end of the flight where the CVR suddenly stopped (emphasis added). [The diagram] End of tape time series, shows the presence of these 0.86 transients which were demonstrated by independent structural and flutter analyses to be quite close to the frequency experienced by a damaged engine mount.”

    Please see Compact Disc enclosed inside back cover for digital demonstration of the “triboelectric” effect of electroacoustical transduction.

    To summarize then, the CVR data which was never researched by the NTSB, hence not available for inclusion in the official report of this accident:

    • There are two shutoff spikes recorded on the CVR, approximately .263 seconds apart, at the end of the recording. There is a highly divergent, whirl flutter type signal immediately prior to the first violent event, detectable both acoustically and analytically.

    • The frequency of vibration of the acoustical event prior to the first spike corresponds to the frequency previously found to be the wing torsion asymmetric vibration mode for the engine in vertical and lateral translation. In other words, the structural frequency detected at the CVR is what would be present had the engine been in the process of separating from the wing.

    • Half of the acoustic signal associated with the four-blade passage was lost at the end of the tape. Simply put, the CVR detected only one engine/propeller combination producing power after the first violent event recorded on the tape. At that point in time, it can be implied that one engine departed the aircraft.

    • The CVR terminated with a 5 G (or greater) load, with no sounds of aircraft impact with the water, due to the loads imparted to the empennage by the right engine as it struck the right horizontal stabilizer.

    • An intermittent amplitude modulation was found on the “silent” track, which occurred over the entire thirty two minutes of the tape. Its frequency correlated with the predicted frequency of a damaged engine mount vibration just prior to and at the onset of the whirl flutter event. Thus, the CVR acted as a latent transducer to not only confirm the whirl flutter event, but also to warn of the existing engine mount damage at least 30 minutes prior to the catastrophic event

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    Sequence of Events, Based on CVR Analysis and Wreckage Documentation

    * Power is reduced on right engine from a “zero thrust” setting to flight idle. This produces the sound of the first gear warning horn “beep”.

    * Whirl flutter is induced in the right engine/propeller due to cracked or failed truss tubes in the right engine mount truss tube assembly. Whirl mode flutter is acoustically evident in the fifth gear warning “beep”. Pre-existing truss tube failures are electroaccousticaily evident in CVR transient signals, throughout 32 minutes of the tape.

    * The right engine and nacelle separate violently from right wing. First significant “spike” on CVR is preceded by a violent acoustical event. Frequency of vibration of this event is exactly that expected if engine were to separate from the wing in flight (see diagram #6).

    * The right engine/propeller translates aft and strikes right and left stabilizer. This causes at least a 5g force upon the tail and the termination of the CVR recording.

    * The right engine completely removes right stabilizer from aircraft. Right propeller impacts left stabilizer, causing the characteristic leading edge damage and bleed air line cuts. Fiberglass nose cone at very top of T-tail is not damaged.

    * As the right horizontal stabilizer is torn from the aircraft, the right elevator “nose down” trim cable is cut, and the “crossover” cable is pulled violently and stretched. The “nose down” cable is then pulled violently in the opposite direction until mechanically fouled in the pulley. The cockpit trim wheel moves well beyond the nose down trim stop.

    * As the right horizontal stabilizer is torn from the aircraft, the tail twists rapidly clockwise (as viewed from above). This twist is “set” into the vertical stabilizer.

    * With the removal of the right horizontal stabilizer, the damage to the left horizontal stabilizer and the extreme nose down trim, the aircraft pitches over violently.

    * Portions of both wings fail instantly in downward bending. The right wing separates at approximately WS 124, at the nacelle. The left wing fails at WS 211.

    * The outboard portion of the right wing, although separated from the fuselage, remains attached (temporarily) to the lower spar cap. This cap then “zippers” out of the bottom of the wing skin, completely bisecting the right main landing gear shock strut.

    * The aircraft is now completely uncontrollable. All normal electrical and hydraulic power is lost. The landing gear is no longer hydraulically held in the up position. The nose gear free falls during descent, into the mechanically locked down position.

    * The aircraft impacts the surface of the ocean, probably nose and left wing root first, in a nearly vertical attitude.
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    Human Performance
    Training Issues

    A detailed review of the CVR reveals that throughout the entire 32 minutes prior to cessation of the recording, the flight is conducted with the highest degree of professionalism. In that time, there are numerous discussions of proper basic flying technique, instrument approach procedures and emergency “drills”, but not even one spurious or casual remark that did not directly pertain to the training environment.

    it is important to remember that this is not a “check” ride for the student, Mr. Lurie. Any evaluation being made by the instructor, Mr. Murphy, at this point in the training scenario are only to identify those obstacles to satisfactory performance that impede the student. Through demonstration, communication and repetition (specific drills), the instructor attempts to correct the problem, not “pass” or “fail” the student. Additionally, there are times when repeating and successfully completing maneuvers that had been difficult for the student in the past instills confidence in the trainee.

    As stated in the FAA’s “Aviation Instructor’s Handbook" [19] "..Those things most often repeated are best remembered. It is the basis of practice and drill”, and “Every time practice occurs, learning continues. The instructor must provide opportunities for students to practice or repeat and must see that this process is directed toward a goal.”

    Captain Murphy is an extremely thorough but demanding instructor. Throughout the flight, he maintains total control of the situation, while allowing the student room to make those minor errors which are so essential for effective training. At no time does the instructor ever let the training scenario go beyond the reasonable limits of safety.

    It is evident that the instructor detects basic weaknesses in the performance of the student, particularly in the areas of “partial panel” instrument flying, single engine work and instrument approach procedures. There are numerous discussions regarding these particular maneuvers recorded on the CVR, and in fact, instructor Murphy has the student repeat these procedures several times in order to strengthen the trainee’s abilities. Murphy does load the student up with a lot of work, but on all occasions he gives the trainee ample time to stabilize one situation (i.e., partial panel) before introducing another (engine failure). Lurie is a “captain” candidate, and as such, has to be able to handle difficult situations such as those presented to him by his instructor.

    One FAA inspector, upon being interviewed after this accident, stated that if an applicant for an ATP or type rating (which Lurie was) had lost his attitude indicator during flight, he would have expected the applicant to be able to fly partial panel without referring to the other attitude
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    indicator[20] Certainly Captain Murphy was aware of this.

    The precise reasons the instructor created the scenario he did are unknown. However, it would appear that it is an attempt, through repetition and exercise, to correct basic student weaknesses. It may also be that because of Captain Murphy’s knowledge of this particular aircraft’s systems, he was attempting to simulate failures that, while seemingly unrelated, may in fact be realistic because of the basic design of the aircraft’s systems. In any case, this particular scenario did not technically introduce “multiple emergencies” as defined by the FAA (see Board report, page []), but did emphasize those areas of student weakness that were of concern to instructor Murphy. Those same areas would ultimately need to be demonstrated to an inspector during Lurie’ s upcoming check ride with the FAA.

    Altitude Alert Chime

    The altitude alert “chime” or tone is recorded on the CVR 8 times in the final 32 minutes. This chime is activated either by the aircraft climbing through or descending away from an assigned (and “set”) altitude by 300 feet, or by resetting the altitude in the alerter control panel. For example, if the aircraft is level at 3,000 feet, and 2,000 feet is then selected in the control panel, the altitude alert chime would sound as the alerter is dialed through the 2700 foot setting towards the 2000 foot setting.

    We can assume that the pilot flying has the responsibility to set the controller, as evidenced by Murphy’s comment on the CVR at 2120:37 “ah, what’d you put in there?”, and Lurie’s remark, “I didn’t, that’s just it, huh”. It is obvious in looking at the entire conversation that both comments refer to the altitude selected in the altitude controller.

    On at least three occasions, the alert chime sounds due to the setting of the controller by one of the pilots. At 2117:39 “...ah, the magic number as I said, is two thousand”. The aircraft at this time is at 2,500 feet. At 2117:46, the chime sounds. While reviewing the approach plate, Lurie sets the 2,000 foot initial altitude into the controller, causing the chime to sound.

    Later, during a different approach, there was a similar situation. Again, at 2141:16, the comment “two thousand after I cross the VOR”. At 2141:44, the altitude alert chime sounds. The aircraft is still at 2,500 feet. At 2142:14 the comment “out of two and a half for two thousand”. The altitude alerter is set prior to, and in anticipation of the descent out of 2,500 feet, which is initiated 30 seconds after the alerter is set.

    Both of these events occur during briefings regarding proper altitude awareness during descent from one intermediate approach altitude to another. Both occur approximately 30 seconds prior to planned aircraft descent.

    The last time the altitude alert chime sounds due to the flying pilot changing the setting is the last chime recorded on the CVR, at 2146:34. Student Lurie has had a tendency to descend out of the
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    initial approach altitude early, as indicated in the prior two approaches flown that evening. Both times he started descent to MDA before completely established on the inbound final approach course of the non precision approach, and both times instructor Murphy “coached” him in the proper approach technique.

    At 2145:30, student Lurie asks, “what altitude I’m still good down to now”. He is anticipating his descent to the MDA and the resetting of the altitude controller. Instructor Murphy replies “ah, two thousand still”. In other words, no descent yet, because the aircraft is still on the initial approach segment. At 2146:27 Lurie again asks, “what altitude am I good down to?” still anticipating the descent. Obviously the student feels that he is close to intercepting the final approach course at this time, probably within the same general 30 second time frame he has used on the previous two approaches. Murphy responds “ah, once you’re established inbound, right, you’re good down to what?” This reinforces the procedure of waiting until interception of the course to initiate descent, but also forces Lurie to think more about the correct approach procedure, and perhaps allowing less concentration on the task at hand, flying the aircraft. In response to the question, the student pauses, then reaches up and selects the new altitude in the controller. This diverted thought/action process distracts him somewhat, and probably results in an unwanted aircraft roll due to asymmetric engine thrust. Murphy immediately responds with some “coaching” to keep the aircraft’s flight regime within safe parameters, and responds, “get the bank” .

    The assumption that the final altitude alert chime is due to the aircraft departing it’s “set” altitude of 2,000 feet in an uncontrolled descent is in error. At no time prior to inflight breakup does the aircraft depart 1,900 feet. This is well documented by the 1,900 foot mode C altitude returns that continue up to the last sweep of the CDR data.

    It must be remembered that the only instrument in the cockpit that is inoperative is the left side (student’s) attitude indicator. Both altimeters, vertical speed indicators and airspeed indicators etc. are functional. Because of the greater number and quicker reacting “cues” available for altitude (i.e., altimeters, vertical speed and airspeed indicators), it is inherently easier to maintain pitch attitude, therefore altitude, than it would be to maintain roll attitude with an inoperative attitude indicator. Therefore, we would expect the student to have more trouble maintaining “wings level” than maintaining altitude.

    And we see this with the instructor’s comment, “get the bank”. Murphy has carefully monitored aircraft altitude throughout the flight - there is no reason to suspect he is not aware of the aircraft altitude at this point, and he makes no comments regarding any altitude deviations near the end of the tape.

    Situational Awareness

    The probable cause of this accident, as determined by the NTSB was partially “...the instructor pilot’s loss of altitude awareness and possible spatial disorientation, which resulted in the loss of control of the airplane at an altitude too low for recovery...

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    There are 32 minutes of recorded conversation on the CVR prior to the cessation of the tape. In that time, instructor Murphy continually comments on the airplane’s altitude, the student’s altitude awareness and the proper aircraft attitude. There are 10 different discussions of descent to MDA prior to final approach course interception, totaling several minutes. There are over 15 different comments and discussions that reveal the instructor’s total command of the “situational awareness” of the flight. ‘You’re eight miles from the beacon...” (2118:17), “Just be one step ahead of the plane. Know you’ve got to get to the VOR (procedural discussion)...” (2121:04), and “..want to be aggressive to get on that course outbound now.” (2142:47), are a few examples.

    There are over 20 different instances of Murphy’s coaching the student regarding altitude and attitude, including “...already at eight hundred feet.” (2115:03), “watch altitude..” (2128:26), “don’t climb” (2114:55), “watch what you’re doing...” (2124:31), “what speed do you want?” (2121:30), “fly the airplane first...” (2114:36), “watch your altitude, altitude, altitude...” (2130:56), “now think what you’re doing, one fifty ...“ (213 1:13). And finally, at the end of the recording, “Stop one thing at a time...” and “get the bank” (2146:44).

    All of these indicate an acute awareness on the part of the instructor, of the attitude and altitude of the aircraft. Obviously in order to state “get the bank”, one of the final comments on the CVR, Murphy has to have referenced his own operating attitude indicator, processed the information, determined that the trainee can provide proper control inputs to keep the aircraft on profile, and has relayed that information to the student.

    At no time during the flight, and specifically in the last few minutes, is there any stress evident in instructor Murphy’s voice. None of the spectroanalysis done indicates any fluctuation or deviation from normal voice patterns, which would be present if there was any indications of “alarm” in his comments. Captain Murphy maintained aircraft attitude and altitude control throughout the flight by reference to his own instruments. There is absolutely no evidence that would indicate that Capt. Murphy would fail to notice any deviation from desired altitude. The termination of the CVR and the breakup of the aircraft was due to an instantaneous, catastrophic event at altitude to which the pilots had no warning and from which they could not recover.

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    Summary Findings

    #1. The flightcrew was qualified and current in accordance with FARs and company policies.

    #2. All training was conducted in a professional manner. The instructor pilot (IP) was concentrating on those areas of weakness for the captain-trainee. specifically “‘partial panel” flying and single engine approaches.

    #3. The Beechcraft 1900 has a well-known history of engine mount truss tube cracking, separation and failure. The original truss assembly has undergone extensive design and manufacturing changes in an attempt to alleviate some of these problems.

    #4. Through electroacoustic transduction, the CVR continually records the transient structural frequencies normally associated with the pre-existing condition of two failed engine truss tubes.

    #5. Very near the end of the flight, an excessive bank angle probably resulted 4ue to trainee distraction and asymmetric thrust. However, because airspeed was well below configuration maneuvering speed, pilot control input could not have resulted in overstress to the airframe.

    #6. Power was reduced on the right engine from a zero thrust setting to the flight idle setting. A whirl mode flutter was induced to the right engine and propeller assembly, due to pre-existing right engine truss tube failures.

    #7. This whirl mode flutter is evident in the CVR spectroanalysis, and is followed immediately by a violent acoustical event and instant cessation of the CVR recording.

    #8. The whirl mode flutter caused a catastrophic failure within the truss mount system. This failure allowed the right engine and nacelle to depart the right wing, evident as a strong acoustical event recorded immediately prior to the first ““spike” on the CVR.

    #9. The right engine struck and removed the right horizontal stabilizer. The propeller probably damaged the left horizontal stabilizer. The aircraft then pitched over violently and instantly. The CVR stopped recording when the 5g limiting switch was tripped by the force of the engine striking the empennage.

    #10. Both wings failed upon pitchover, and the aircraft was no longer controllable. After descending 1900 feet, the fuselage of the aircraft impacted the surface of the ocean in a nearly vertical attitude.

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    The complete set of diagrams, photos and reports are not reproduced at this web site, except one diagramdepicting the tail damage (Figure 2) is attached to this file. Also,. a paper describing Stearman's work with the voice recorder, which includes audio and visual illustrations can be accessed at


    [1] National Transportation Safety Board, “Loss of Control, Business Express, Inc., Beechcraft 1900C N811BE Near Block Island Rhode Island, December 28, 1991”, PB93-910405, NTSB/AAR-93/O1/SUM, April 27, 1993
    [2] ALPA has included approximately 60 photographs as part of this exhibit. An additional 350 photographs are available for review upon request.
    [3] Bowers, David F., Ph.D., Packer Engineering, Inc., “MG - Business Express, Inc., SN UB-49 Beech 1900C”, March 1, 1993
    [4] Cauble, Robert F., Associated Data Resource, “Boston ARTCC & Ocean Terminal Radar Control, Recorded Radar Data Report, Radar Data Study, Data List, Plots”, 1994
    [5] “Beechcraft 1900 Series Maintenance Manual”, P/N 1 -0021-7a26, Beech Aircraft Corporation, 1988
    [6] "Staff Study, Model 1900/1900C engine Truss Fatigue Cracking". Beech Aircraft Corporation, Oct 19, 1990
    [7] Stearman, Dr. Ronald, P.E., Buschow, Monte, Kane, Kevin, “The Beech Aircraft Corporation Model 1900 Airliner Engine Truss: A Study in Reliability Analysis & Aviation Safety”, June 3, 1995
    [8] Complete memorandum and cover letter are included in attachment C.
    [9] Complete SDR list compiled by Air Data Research May 20, 1997 available upon request
    [10] MSC/Nastran -“ Handbook for Aeroelastic Analysis” Vol. 3, Version 68 Mac Neal-Schwindler, 1995, USA. This is a commercially available finite element modeling code. This study used the whirl flutter program resident in the aeroelasticity option.
    [11] Brock, B., “Personal Communication Concerning the Modified XC 142 Whirl Flutter Software Code”, 1966. A refinement of this system was used, using a basic whirl flutter software code developed by Voight Aeronautics
    [12] Stearman, Dr. Ronald, P.E. , Schulze, Glen H., Rohre, Stuart M., “Aircraft Damage Detection from Acoustic and Noise Impressed Signals Found by a Cockpit Voice Recorder”, 1997, Institute of Noise Control Engineering.
    [13] McSwain Richard H., Ph.D., P.E., Investigation Conclusions, Materials Engineering Investigation. The original reports includes approx. 410 photographs and supporting microscopic examination documentation, not included in this petition but available upon request.
    [14] Hammill Donald F., “Expert Report”, 1995
    [15] Pratt & Whitney Canada, Service Investigative Report, “Business Express Beech 1900C NS11BE, Block Island, Rhode Island”, Report #TL-852, August, 1992
    [16] Zwillenberg, Peter, Report 1900JE361 D, page 80, Beech Aircraft Corporation, September 30 1983, as quoted in Attachment D, page 47
    [17] Reed, W. III, “A Review of Propeller-Rotor Whirl Flutter”, NASA TR R-264, 1967
    [18] Schulze, Glen H., “Cockpit Voice Recorder Tape Erasure Gap Study & Signal Level Inventory and Study”, Data Acquisition Systems, Littleton, CO. February 26, 1996
    [19] "Aviation Instructor's Handbook", US Department of Transportation, Federal Aviation Administration, 1977, page 3
    [20] NTSB report of subject accident, NTSB/AAR-93/O 1/SUM, April 27, 1993, page 18