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This report describes results of an alternative investigation resulting in the following 1990 Petition for Reconsideration of an original investigation report by a government agency. This document is noteworthy for several reasons, including good use of direct observations of their experiences which crew members provided to investigators. It also provides interesting arrays of the data. From a research perspective, it is also interesting to compare the contents of this petition with the kinds of alternative analyses posted elsewhere, such as those offered at the Ladkin and GAAG sites. It might be viewed in the context of Johnson's paper Improving the Presentation of Accident Reports over the World Wide Web

The original report is referenced at the NTSB web site, but is not reproduced there. A copy can be acquired from the National Technical Information Service by referencing the NTSB report title NTSB Aircraft Accident Report NTSB-AAR-81-8. I am trying to locate a copy o fht report and will post it if I can get my hands on one.

If the Board responded to this petition, and another 1991 Petition by TWA, the response are not yet available. The NTSB's December 1983 response to a 1983 petition by ALPA and the NTSB's May 4 1995 response to a 1991 Petition for Reconsideration by the Pilot in the same case are posted here for review after reading this report. Comments about the reports and the underlying investigation processes are invited for the Forum.

This document was contributed by Leigh D. Johnson, who helped format it for publication here. The responses will continue to be pursued.



Use your Browser FIND menu to go to desired text in this table of contents.
(Aclickable INDEX of contents is found at the end of the footnotes.)

 

Petition for Reconsideration of Probable Cause

(High Altitude Yaw--Roll Over -- Vertical Dive accident)

TWA Flight 841
Boeing 727-100
Saginaw, Michigan
April 4, 1979


Prepared by the

Air Line Pilot's Association

 

 

Proposed changes to the NTSB's

Aircraft Accident Report AAR-81-8

are detailed in the pages herein.

9/4/90 Table of Contents page i

Illustration, NTSB's Erroneous Assumptions iv NTSB Form 1765.2, Technical Report Documentation 2
Section 1. Factual Information 2
Section 1.6 -- Aircraft Information 2
Section 1.11 -- Flight Recorders 3
Section 1.12 -- Wreckage and Impact Information 5
Illustration, Component Failure Diagram 7
Illustration, Post Accident Inspection 8
Illustration, Maintenance Records 9
Section 1.16.1 -- Boeing Company Tests 12
Section 1.16.2 -- Flight Simulator Tests 12
Section 1.16.3 -- Heading Gyro Tests 15
Section 1.16.4 -- Flight Tests 15
Section 1.16.5, TWA Tests Aboard N840TW 23
AFFIDAVIT 24
Section 1.17.1 -- Under B-727 Flap System 26
Section 1.17.2 -- History of B727 Leading Edge Slat Problems 26
Section 1.17.3 -- Aircraft Performance 26
Section 1.17.4 -- No. 7 Leading Edge Slat Operation 27
Section 1.17.5,-- Loads On A Slat 28

Loads on an Extended Slat 28
Loads on a Retracted Slat 30
Section 1.17.6, B727 Rudder Control Systems 31
Section 1.18 -- Useful or Effective Investigative Techniques 36
Section 1.19 -- Efforts to Revise Investigative Errors 37
Section 2. Analysis 38

Section 2.3 -- The Aircraft 38
Section 2.4 -- Direct Evidence 39
Section 2.4.1 -- The Trail Of Debris 40
Illustration, Trajectory Analysis 41
Section 2.4.2 -- Gear Extension Damage and Hydraulic System Failure 42
Illustration, Landing Gear 44
Illustration, Damaged Gear 45
Illustration, Recovery--Failure Sequence 46
Illustration, Continued Failure Sequence 47
Illustration, The Recovery and Trail of Debris 48
Section 2.4.3 -- Limited Strength of an Extended Slat 49
Illustration, Limited Strength of an Extended Slat 51
Section 2.4.4, -- Evidence of Sustained Sideslip 52
Section 2.4.5 -- Pre-existing Leakage From A Worn Actuator For #7 Slat 53
Section 2.4.6 -- Extension of the No. 7 Leading Edge Slat 54
Section 2.4.7 -- Sequence of Separation Of Parts During The Recovery 60
Illustration, Profile View of Recovery Pull Up 62
Illustration, Overlay of Aircraft Track and Parts Trajectory 63
Section 2.5 -- The Initial Upset and Loss of Aircraft Control 65
Section 2.5.1 -- Vibration, Yaw,and Roll 65
Illustration, Lateral and Directional Flight Controls 67
Illustration, Initial Right Rolling Moment 68
Illustration, Yaw Damper Interaction 71
Illustration, Negative Sideslip Angle Caused By Yaw 72
Illustration, Vertical Stabilizer and Rudders 73
Illustration, Sweepback and Dihedral Effects 74
Illustration, Left Wing Flight Spoilers Extend -- Spoiler Buzz 75
Illustration, Yaw Damper Fail Flag 77
Illustration, Rudder Position Indicator 79
Section 2.5.2 -- Freeplay and Flutter of the Right Outboard Aileron 80

Vibration 83
Section 2.5.3 -- Autopilot Roll Channel 86
Section 2.6 -- Fault Analysis of Slat Retract Locking Mechanism 89
Possible Failure of Retract Lock Keys to Engage 91
Possible Failure or Absence of the Retract Lock Piston Spring 92
Possible Failed Seal Around the Sensing Pin for the Retract Switch 94
Section 2.7 -- Fault Analysis of Slat Retract Lock Indicating Switch 95
Section 2.8 -- Investigative Errors Sparked Erroneous Findings 96
Synthesized Pitch and Roll History 97
Evidence Selection and Analysis 98
Sideslip Effect on Lateral and Directional Stability 101

Section 3 --
Conclusions 104
Section 3.2 --
Probable Cause 108

Section 4 -- Recommendation 109

NTSB Record Keeping 109
Table, Correlation of FDR Time with Evidence 112

 

Summary of claimed errors.


Petition for Reconsideration

of Probable Cause

TWA Flight 841

Boeing 727-100

Saginaw, Michigan

April 4, 1979

The NTSB erroneously assumed that an extended slat caused the upset of TWA 841.

NTSB Reg §845.41(a) states,

Petitions for reconsideration or modification of the Board’s findings and determination of probable cause . . . will be entertained only if based on the discovery of new evidence or on a showing that the Board’s findings are erroneous.

In this case, the April 1979 B727 high altitude yaw -- roll over -- vertical dive accident, no new matter need be presented to show that the Board’s findings are not consistent with the evidence already included in the docket. However, some new information is included, in the form of an affidavit regarding a 1977 flight test at TWA and the resulting flight control problems encountered.

The Board should review the historical records of the other yaw, roll-over mishaps.

The current NTSB-AAR-81-8 contains numerous erroneous assumptions labeled as “findings,” “conclusions,” or “determinations.” The Safety Board regarded circumstantial evidence, created after the accident, as superior to direct evidence. The manufacturer of the accident aircraft had developed this circumstantial evidence at the their own facilities. The NTSB should have regarded such evidence with appropriate skepticism.

The following sections offer the Board specific directions for rewriting the NTSB Report. The Board erred by assuming that an extended slat had caused the upset of TWA 841. That erroneous assumption tainted the investigation. In several instances, the Board failed to document factual information properly. The Board employed incomplete and improper analysis methods in their effort to substantiate that erroneous assumption. The following pages include criticism on many of the items covered in the NTSB Report. The following pages show proposed corrections to many paragraphs in the report. Portions of the Board’s Analysis, and Conclusions, were not pertinent to the failure sequence suffered by the accident aircraft. In many cases, the Board should discard whole paragraphs from their report. Included are several proposed changes to the structure of the Board’s Report so they can present the evidence in a logical fashion.

The NTSB Report -- Identified Oversights, or Errors

NTSB Form 1765.2, Technical Report Documentation page, pg i, included statements that reflected the analysis errors that attributed the loss of control to an extended slat. Also, item 17, “Key Words,” should probably include the labels: yaw, roll over, vertical dive, yaw damper failure, and rudder system failure, failure of an outboard aileron, and cockpit voice recorder malfunction.

Section 1. Factual Information

Page 4, underAIRCRAFT INFORMATION , Section 1.6describes the 1Mar79 “C” Check. The NTSB did not include reference to the hydraulic leakage in the area of the lower rudder actuator. The NTSB also failed to mention that mechanics had repainted skydrol stains on the right wing upper surface. After the 4Apr79 accident, investigators found evidence of skydrol bathing aft of the #7 slat7 slat. The Maintenance Records Group included these items in their Report, found in the NTSB docket.

Page 6, underFLIGHT RECORDERS , Section 1.11. In describing the Flight Data Recorder (FDR) output, this paragraph of the NTSB’s Report minimized the two errors that affected the heading trace. The combined effects of directional gyro gimbal error, displayed as yaw reversal, and stylus anomalies, displayed as shifts backward in time, made the heading trace virtually unusable for purposes of dynamic analysis. This combination of errors affected the heading trace at the crucial interval during the initial yaw-roll upset.

Page 6, underFlight Recorders , Section 1.11, states:

Tests of the CVR in the aircraft revealed no discrepancies in the CVR’s electrical and recording systems.

The Cockpit Voice Recorder (CVR) Group Report, included in the docket, did not substantiate that statement regarding onboard testing of the CVR. Review of the NTSB docket revealsNOdocumentation of any sort of CVR system check-out or even a bench check of the recorder unit. Airline employees extracted the CVR recorder unit from the accident aircraft within hours of the event. FAA Form 3112, Inspection and Surveillance Record (as completed by Roger V. Gordon, Jr., dated 4-5-79), disclosed that he confronted the TWA Station Manager (a Mr. Frank Cook).

He stated, he had been instructed by his Supervisors in Kansas City not to give the recorders to FAA or anyone else, the recorders were locked in the trunk of his car.

The electrical connections that existed at the time that the accident aircraft landed at DTW were thereby lost, compromising the integrity of any system check-out that might have been attempted later.

After landing at DTW the normal electrical power change-over from engine generators to APU generator occurred shortly after landing. The flight engineer started the APU, then switched the aircraft electrical loads to the APU generator. This power transfer occurred just prior to engine shutdown while the pilot parked the aircraft on the taxiway near the runway. It may have been this electrical “spike” that either: 1) caused a functioning CVR to initiate an erasure; or 2) awakened a non-functioning recorder causing it to begin recording until mechanics removed electrical power 9 minutes later. A normal CVR would note the power change-over by imprinting a 600 Hz tone on the tape. [1]A fault could cause the CVR to activate the bulk erase field instead of the power change-over tone.

There is no annunciation of CVR faults displayed in the cockpit. The CVR could cease recording for extended intervals and still the system would provide no annunciation to the pilots.

One other factor not mentioned in the NTSB Report was the installation of the recorder unit (with exposed switches) in the aft baggage compartment. This CVR installation was wired to a squat switch in the gear well, which provided the interlock that enabled the CVR erase function. This particular CVR installation, combined with damaged wiring in the gear well, may have permitted a discrepant “erase” function to be initiated during the uncontrollable segment, or during the landing roll out.

A fault may have caused the bulk erase coil to broadcast the 400 Hz field over the tape reel during most of the life the that recorder unit. There is no record of any sort of post accident fault identification attempted. There is no record of tests of that particular CVR system (as installed on aircraft 7840) nor of CVR systems on other aircraft in the 727 fleet.

Investigators had found inoperative CVRs on other accident aircraft. Following the UAL DC-8 accident at SLC, CVR anomalies turned-up fleet wide. The airline accomplished fleet wide CVR fault checks at UAL. These fault checks were said to have demonstrated that the self test features of the CVR installed on their DC-8 aircraft were unreliable, and did not reflect faults in the CVR system. A careful investigator would have considered thepossibility of an intermittently malfunctioning CVR in the case of the TWA 841 accident.

Section 1.12 -- Wreckage and Impact Information . The Board inaccurately identified the damaged flap track and flap track “canoe” fairing, on AAR page 7, paragraphs 2 and 3. The B727 right inboard flap has two flap tracks. The inboard (I/B) track on the I/B flap can be clearly termed the #5 flap track. It was this #5 flap track, closest to the right landing gear strut, that investigators found damaged. The mid section of the #5 flap track “canoe” fairing had separated from the aircraft. Investigators later located that “canoe” fairing in the trail of debris found in a field north of Saginaw.

The debris found on the surface related to a failure sequence. Because of inexact identification of the damaged components, the NTSB was ambiguous about the part of the inboard flap that sustained the damage. This confusion, about the damaged flap track and fairing, also existed in the manufacturer’s report (page A-5 seemingly contradicts page B-30) to the NTSB. The manufacturer used inconsistent or conflicting labels throughout that document.

The Board erroneously identified the missing spoiler panel.

The No. 10 flight spoiler panel . . . was missing. [2][NTSB’s AAR page 7, second paragraph.]

That statement mistakenly suggested that the most outboard spoiler on the right wing -- the spoiler panel aft of the #7 slat -- had separated from the aircraft. That outboard flight spoiler is powered by Hydraulic System “A”, which ruptured during the incident.

Such erroneous component identification probably contributed to the NTSB’s inability to correctly re-construct the sequence of failures suffered by the accident aircraft. Contrary to the above quotation from the NTSB’s AAR, the #10 Flight Spoiler remained attached to the wing.

The B727 upper wing surfaces have twoground spoilers, and fiveflight spoilers on each wing. These flight spoilers are outboard of the ground spoilers. Flight spoilers are identified by number from left to right. Flight spoilers #6 through #10 are located on the right wing. The NTSB Report, page 15, showed a diagram of the B727 flight controls, with the two most inboard spoiler panels properly labeled as ground spoilers.

Photographic evidence [3]showed clearly that the #10 flight spoiler remained attached to the wing. The photographs showed that themost inboard flight spoiler on the right wing, the #6 flight spoiler, was missing. The photographs of the upper surface of the right wing showed only one missing spoiler panel. That panel was missing from the position located above the outboard track of the inboard flap. Only when this missing component is correctly identified, can it be logically fit into the failure sequence sustained by the accident aircraft.

The spoiler panel that ripped from the accident aircraft was the #6Flight Spoiler , the most inboard flight spoiler panel on the right wing. The separated portion of that spoiler panel was found with the trail of debris, in the field north of Saginaw. The #6Flight Spoiler is powered by Hydraulic System “B”, which remained intact throughout the flight.

The NTSB report should have presented an illustration of the B727-100 aircraft, with documented component damage clearly labeled.

Intentionally left blank.

Yaw Damper Fail Flag, Landing Gear Damage, aircraft damage illustration, Rudder Boost Packages, yaw damper transfer valves, aileron damage, slat #7, flap track “canoe” fairing, spoiler panel damage illustration, hydraulic line damage,

Component Failure Diagram

Intentionally left blank.

skydrol bathing, wing surface, slat #7, Boeing Ops Manual Bulletin, aileron damage, hydraulic leakage ,

Post Accident Inspection

of Right Wing Upper Surface

Intentionally left blank.

hydraulic leakage, slat #7 pre-existing damage, skydrol damage, rudder actuator leakage,

Maintenance Records

sfrom prior “C” Check

Under Wreckage and Impact Information , on AAR page 7, paragraph 3; the Board’s description of the orientation of the debris disagrees with data presented in the report from the manufacturer, page A-10 Figure 4. The Board must carefully describe the distance and direction between aircraft components found on the ground. This trail of debris is the best evidence available in this accident and can be used to reveal the sequence of some failures sustained by the accident aircraft.

The NTSB report should have shown the trail of debris properly mapped with accurate scale distance and direction between components. In the later sections of their report the Board should fully analyze such a map of the trail of debris and explain the sequence of failures.

Under Wreckage and Impact Information, on AAR page 7, the last paragraph should clearly state that investigators did not check the annunciation for the #7 slat retract locking mechanism. The current wording of the paragraph may mislead the reader to think that the annunciation for the #7 slat retract lock was found to function properly. (See admission on page 24 of the NTSB’s AAR, mid-page: “the #7 actuating and indicating system could not be checked . . . .”)

Under Wreckage and Impact Information , on AAR page 8, the second paragraph misleads the reader to think that investigators had examined the #7 Slat Retract Lock Indicating Switch. Records in the NTSB docket show that post accident examination of the remnant of the #7 slat actuator focused on the fracture to the outer cylinder barrel of the actuator. None of the records in the NTSB docket reflect any functional test nor tear down analysis of the Retract Lock Indicating Switch (illustrated on page 17 of the NTSB Report).

In the docket a letter dated February 20, 1981 from L. D. Kampschror, NTSB Investigator in Charge, to Capt. J. A. McIntyre, ALPA, stated:

The actuator that was on N840TW had no spring in its switch mechanism which imposed a load or force on the locking keys. [4]

In a letter dated March 5, 1981, Capt. McIntyre responded:

If this statement means the specific actuator remnant for #7 slat on N840TW was found to have no spring force in its switch mechanism, you are revealing a defect in this actuator not hitherto disclosed. . . . Switches measured in TWA’s shop indicated a nominal force of 13 pounds required to actuate the switch.

In their report the Board should have included the information provided in the quotations above. Such a Slat Retract Lock Indicating Switch, lacking any internal spring force, would be unable to activate the amber slat in-transit annunciation in the cockpit. This amber annunciation would have been the only clue to the pilots that the mechanical slat actuator retract lock had not engaged after slat retraction. Under normal conditions with the slats retracted, the hydraulic pressure would act against the retract side of the actuator piston. However, various forces could pull the #7 slat from the retracted position following a failure of Hydraulic System “A”. This extension would result if the mechanical slat actuator lock were not engaged and the slat was subjected to a combination of tensile loads (G forces, vibration, and aerodynamic forces).

Section 1.16.1 -- Under BOEING COMPANY TESTS ,on AAR page 9; the Boeing OMB 75-7[5]is referenced. This paragraph should also state that the OMB 75-7 outlined specific conditions which would permit a leading edge slat to be pulled from the retracted position (Mach _ .8M, with failure of hydraulics to the slat actuator).

Section 1.16.2 -- Under FLIGHT SIMULATOR TESTS ,on AAR pages 9 and 10, the NTSB’s report writers included erroneous information provided by the manufacturer. The first paragraph underFlight Simulator Tests, on pages 9 states that testing was conducted in a B-727-200 simulator. In the footnote (note 4) the NTSB repeated false information found on page C-4 in the Boeing Report:

The only difference aerodynamically between the two airplanes is a 120_ difference in body length.

First the Boeing Report, and then the NTSB Report, incorrectly stated the magnitude of the difference in body length between the (short body) accident aircraft and the B-727-200 modeled in the simulator tests. In fact, the difference in body length is twice that magnitude (body length of the 727-100 is 116 ft 2 in, the 727-200 is 136 ft 2 in; as published in the manufacturer’s document D6-1420).

Furthermore, the false information provided in the Boeing Report misled the NTSB to neglect other major “stability and control” differences between the B727-100 and -200 models. Such differences include “tail volume,” yawing moments due to rudder, sideslip generated by rudder, yawing moment due to sideslip, yawing moment of inertia, and pitching moment of inertia. The manufacturer’s attempt at post-accident simulation thus suffered from inaccurate modeling, having neglected such differences between their simulator and the accident aircraft. The 118 simulation trials also neglected the effects of the up-float of the right outboard aileron. Their B727-200 simulator could not duplicate the tightly coupled high altitude control characteristics of the accident aircraft. Limitations of the simulator prevented the investigators from exploring possible faults of the lateral and directional control systems that interacted to induce the dynamic yawing, rolling, and pitching upset maneuver of the accident aircraft.

In the first paragraph underFLIGHT SIMULATOR TESTS , on AAR page 9, the Board summarized the simulator testing. The Board should state that the investigators prearranged test conditions, intending that test results would support their assumption that the extension of the #7 slat was a cause of the upset. This paragraph should restate an important admission included in the manufacturer’s report, on page C-1:

It was also determined that the extension of slat #7 at the flight condition of N840TW prior to the incident can be easily controlled with prompt corrective action by the pilot.[Emphasis added.]

That statement suggests that the initial upset of the accident aircraft was NOT induced by an extended slat.

The second paragraph underFlight Simulator Tests , on AAR page 9, should state clearly the test conditions used to simulate uncommanded rudder deflection. The manufacturer, on page C-6 of the Boeing Report, described the method used for simulating yaw damper hardovers during the piloted simulator trials:

. . . yaw damper hardovers were not directly simulated . . .[Emphasis added.]

Regarding yaw upsets, the Board should refer to the results from Project Race, and the CAB’s investigation of the AA Flight One accident.[6]

Project Race was a program of flight tests originated by the FAA to shed light on the cause of the AA Flight One accident. NASA, AA, Boeing and the CAB participated in the tests. The tests measured the response of the airplane to the effects of slips and skids, and malfunctions of the rudder control system. In their report of January 1963, the CAB included a message meant for future accident investigators:

Project Race provided the Board with much information that will prove helpful in future accident investigations. [7]

During the investigation of the AA Flight One accident, the CAB concluded that the manufacturer’s testing of rudder upsets was inadequate.

The tests are obviously planned maneuvers under which conditions the pilot is not confronted with the necessity of analyzing the malfunction, deciding what corrective action he will take, and experimenting to produce the desired results. . . . It is unreasonable to assume that . . . the pilot, confronted with an unexpected roll, would start corrective action as soon and to the extent characteristic of planned flight tests.

The above is borne out by recorded instances of yaw damper malfunction or mismanagement. In all instances the crew was late in recognizing the yaw damper as being the source of the problem and were slow in initiating corrective action. In some cases, even after initiation of corrective action the dangerously steep banked attitudes increased and persisted well beyond flight test values before recovery was effected. In some instances of yaw damper mismanagement the crew never properly analyzed the difficulty and the flights were completed after application of additional lateral control . . . There are some instances wherein the crew took advantage of additional lateral control capabilities, recovered to level flight, analyzed the difficulty, and then disengaged the offending yaw damper. [8]

The NTSB should reiterate those findings of the CAB. The CAB found that pilots had difficulty in accurately identifying the source of an upset induced by a malfunctioning yaw damper. The CAB focused on a B707 type of design: a single rudder, with only one yaw damper. It is much more difficult to identify the source of a rolling motion, caused by yaw, when operating the B727-100 type aircraft at high cruise altitude and high gross weight combinations.

The third paragraph underFLIGHT SIMULATOR TESTS , on AAR page 9, should mention that investigators did not consider the effects of ratcheting of the aileron control wheel. Such ratcheting would reduce the lateral control margin.

The Board was unaware of the results of the 1977 flight tests done at TWA. That testing resulted in an incident characterized by near loss of control. That incident occurred aboard aircraft N840TW. The 1977 test conditions explored the complex problem of an autopilot false disconnect and the ratcheting of the aileron control wheel. In a revised NTSB Report the Board should adopt two proposed new sub sections. One proposed new sub section, 1.16.5, should describe the TWA tests aboard N840TW. The Board should analyze the 1977 TWA test and compare the results with the accident. We include an analysis of this lateral control anomaly in a proposed new sub section2.5.3,titled “Autopilot Roll Channel.”

Section 1.16.3 -- HEADING GYRO TESTS ,on AAR page 11. The combined effects of gyro gimbal error and recorder stylus shifts (due to worn recorder mechanisms), made the FDR heading trace unreliable at crucial moments. The Board should limit their use of the specious heading data and acknowledge the weakness of conclusions based on that heading data.

Section 1.16.4 -- See the first paragraph underFLIGHT TESTS , AAR page 12. The Board should honestly state the nature of the test conditions for the October 2, 1980 flight test. The nature of the flight test was recorded in the transcript of the post-flight conference:

The purpose of the test . . . is to provide a signature on a flight data recorder for several trailing edge flap and leading edge slat configurations. [9]

An assumption guided investigators as they planned the flight test. The intent of the investigators was to design test conditions that would yield useful results. The Board assumed that the extension of trailing edge flaps, and leading edge slats, had occurred immediately prior to the upset of the accident aircraft. Consequently, the flight test results are of little value in identifying any alternative failure sequence.

Section 1.16.4 -- Under FLIGHT TESTS , second to the last paragraph on AAR page 12, the statement regarding airspeed decrease is misleading. Following extension of flaps to 2 degrees, the Flight Test Instrumentation (FTI) airspeed trace showed a steady decrease, at a slow rate. The test crew sustained the condition for 39.2 seconds during which the decay of airspeed continued. Had the condition been allowed to continue, the rate of deceleration may have increased as the autopilot altitude hold function commanded increased nose up pitch

Section 1.16.4 -- Under FLIGHT TESTS , see the last sentence in the last paragraph on AAR page 12. The data did not substantiate the assertion that the airspeed increased as a result of flap retraction only. On the contrary, the test director stated during the debrief, “It’s so terribly painful at 39000 feet to accelerate back to the test speed.” (See page 83-1-D7, of D6-44357.) The investigators had pre-defined no test condition with a stabilized interval at TE flaps 5, followed by retraction of TE flaps to 2. An increase in thrust most likely yielded the increase in airspeed. The test aircraft, E209, was equipped with JT8D-15 engines. [10]These engines provided the test aircraft with a thrust advantage over the accident aircraft, which had lower powered JT8D-7 engines.

Section 1.16.4 -- Under FLIGHT TESTS ,the three paragraphs on AAR page 13 related to a single condition from the flight test. During that test condition, the test crew extended leading edge slats 2-3 and 6-7 while at FL390. The Board accentuated data from that one prearranged test condition. The Board needed to support their contention that, just prior to the upset, the accident aircraft had experienced such a slat extension.

The first paragraph on AAR page 13 described the test condition and results. The Board listed the rate of airspeed decrease as 0.50 Kts/sec following the extension of slats 2-3 and 6-7. However, on examination of the test data included in the NTSB docket, the reader finds that the rates of airspeed decay, following extension of slats 2-3 and 6-7, varied over time. Data from the two test recorders also differed. The more accurate FTI airspeed trace decayed at 0.3 Kts/sec.

This comparison of airspeed decay is discussed later in the proposed new analysis section, 2.8 -- “Investigative Errors Sparked Erroneous Findings.”

The investigators focused on the results of that single condition from the flight test. In its endeavor to link that single condition from the test flight with the initial upset interval of the accident aircraft, the Board proposed an inconclusive comparison. In the second paragraph on AAR page 13, the Board cited a similarity of the recorded airspeed decay and a similarity of the recorded frequency of vibration.

A human fingerprint is unique. However, in flight dynamics the airspeed is affected by several variables; one rate of airspeed decay might be a common response to several conditions involving various flight control deflections. In the case of the accident aircraft, the airspeed decay was most likely due to drag increments caused by other factors such as sideslip and extension of flight spoilers on the left wing as the aileron-spoiler mixer responded to autopilot roll commands.

***********

A forensic test should, as a matter of common sense, satisfy three criteria: the underlying scientific theory must be considered valid by the scientific community; the technique itself must be known to be reliable; and the technique must be shown to have been properly applied in the particular case. [11]

***********

The Board’s comparison of the normal acceleration traces (in the second paragraph on AAR page 13) did not meet the above criteria, and did not qualify as sound forensic technique. There is a limited frequency response of the acceleration channel of such flight data recorders. Thus, unrelated vibrations of unequal frequency beyond the response limit of the recorders would appear on recorder foils as the same vibration frequency. The FDR foil G-trace from the accident aircraft indicated vibration onset at the start of the upset. The increasing amplitude of that vibration signature suggests that spoiler buzz and the flutter of the outboard right aileron were main vibration components. The spoiler buzz resulted from the rising flight spoilers on the left wing. Flutter of the right outboard aileron may have resulted from the dynamic loads on that free floating aileron panel. However, numerous other conditions involving buffet or flutter could have produced vibration of a frequency beyond the response limit of the recorder. The accident FDR would have recorded such vibration at that frequency response limit.

The NTSB’s comparison between recorded data from the flight test and recorded data from the accident aircraft proved inconclusive.

In an effort to explain these weaknesses, the Board included false information in their report. The third paragraph on page 13 of the NTSB report stated:

It was determined that during the flight tests . . . a test switch . . . in the DADC . . . had been left in the test (HOLD) position.

The source document for that information was a letter (dated May 11, 1981, included in the NTSB docket) from Boeing’s H.P. Hogue to NTSB’s Kampschror. A close examination of that letter shows that the manufacturer never actually observed the switch in that HOLD position. Prior to the flight test, proper verification of configuration of crucial test components should have been planned by the test engineers. Normally, the test records would have included specific documentation of component configuration variables. However, the manufacturer presented no such documentation of the position of that DADC test switch for the October 1980 flight test.

During the flight test there proved to be a periodic, undamped oscillation, of the FDR g-trace that developed after leading edge slats 2-3 and 6-7 extended. Such an oscillation was not seen on the FDR g-trace of the accident aircraft. The manufacturer was unable to explain this dissimilarity. Then in an effort to correlate this dissimilar result from the test, the manufacturer made an assumption about the configuration of the test aircraft. These were the actual words reported by the manufacturer to the NTSB:

Therecan be little doubt that this switch was in the test position the day E209 was flown. [12]

The manufacturer’s use of such an assumption in test analysis should have provoked the skepticism of the NTSB investigators. Suspicion should have prompted the Board to ask: What other variables were mismanaged during this one crucial test condition? In the statement shown in their official Aircraft Accident Report, the NTSB altered an assumption made by the manufacturer and improperly reported it as fact.

The three paragraphs on page 13 of the NTSB’s AAR are not pertinent to the accident. Analysis shows that the #7 slat on the accident aircraft remained in the more secure retracted position until late in the dive. Furthermore, the eye witnesses, the pilots, testified that they deployed neither trailing edge flaps nor slats prior to the upset at FL390. The passengers aboard the accident aircraft didNOThear the growl of the flap motors prior to the upset. There were numerous passengers seated in the cabin just above those flap motors. Passengers would have felt and heard several seconds of disturbance while the flap motors operated. The T.E. flaps must extend until near the “two degrees” position before they actuate any slats. Direct evidence from the accident refuted the NTSB assumption regarding slat extension prior to the upset. The NTSB ignored this direct evidence. The results from the flight test ranked only as circumstantial evidence, created after the accident at the manufacturer’s facilities. The NTSB erred in its use of such circumstantial evidence to support an erroneous hypothesis.

Section 1.16.4 -- UnderFLIGHT TESTS ,the top paragraph on AAR page 14 improperly described a flight test condition in which the pilot pushed the right rudder pedal to full travel. The exact words printed in the NTSB report are:

the pilot deflected the rudder fully to the right . . . .

In fact, during the test condition, the FTI shows the exact rudder displacement to have been: upper -3.5 degrees, and lower -4.5 to -5.0 degrees. These angles represented the full deflection available from normal rudder actuators. An accident analyst needs to understand that at FL390 cruise conditions, due to the limitations of the hydraulic pressure acting against aerodynamic resistance, the maximum rudder deflection available from the pilots’ pedals is no greater than the deflection capability of a discrepant yaw damper signal.

This particular condition of the flight test may be more pertinent to this accident than any of the test conditions devoted to flap or slat extension. Had the investigators not fixated on the slat that separated during the accident, they might have more thoroughly tested several possible failures involving sideslip induced roll. Unfortunately, in the results from this most pertinent test condition the manufacturer clipped the FTI data plots. The manufacturer did not show the Board any information that described the set-up of the sideslip. That set-up interval would show the response of the aircraft (yaw, sideslip, bank, and pitch) to the increasing displacement of the rudder surfaces. The NTSB should have noticed this data editing and requested a more complete time-history plot. Working with the FDR plot included for the test Condition No. 1.22.013.025.1, an investigator would have found that the sideslip condition was first initiated about four minutes prior. Therefore, a careful investigator would have ordered a more complete time-history plot to include FTI (IRIG) coordinated time 13:38:30 through 13:44:0. Such a “data request” should have ordered time history plots of other parameters in addition to those parameters presented in the original data plots. Such additional parameters include: rudder pedal position (or rudder pedal force), yaw rate, and sideslip angle. These parameters were probably available from the flight test recording system but not plotted on the graphs provided by the manufacturer’s flight test section.

This most pertinent condition of the test plan (test Condition No. 1.22.013.025.1, NTSB #7) began with the aircraft in a nose-right sideslip (a negative sideslip angle). Then the pilot disengaged the autopilot, with the rudder pedal still displaced. The aircraft rolled from 5 degrees left bank through 33 degrees of right bank in 4.4 seconds. The test pilots terminated the condition after 4.4 seconds by releasing the rudder pedal and using left wheel throw, while the aircraft continued to roll right to 41 degrees right bank. The maximum roll rate recorded on FTI was 13 degrees per second; averaging about 10 degrees per second. The FTI heading trace went from 349 degrees, through 352.5 degrees, during the 4.4 second test condition; with the heading stabilizing at 355 degrees 1.5 seconds later. This FTI heading data had been generated by an Inertial Navigation System aboard the test aircraft, and was independent of any alleged gimbal error that may have affected the #1 DG (FDR Heading trace) of the accident aircraft. Following the official test condition, the pilots used 35 to 45 degrees of left (CCW) throw of the aileron wheel to bring the left wing down to level. This generated a peak roll rate of 23 degrees per second. After the pilot released the rudder, the test aircraft rolled through wings level (zero bank angle) after another 4.1 seconds.

The manufacturer clipped the data presentation for the period during which the pilot established the sideslip (test data was unavailable in the docket). Following the 4.4 second test condition, the data is of little use in the analysis because the test crew rolled the aircraft into a sustained left bank while climbing.

The FTI sampled the positions of the upper and lower rudders during the flight test. The manufacturer presented some test data, with the deflection angle of each rudder plotted in units of degrees. These plots show that during the sustained sideslip, the upper rudder was at -3.5 degrees (the minus sign signifies TE right) and the lower rudder was -5.0 degrees. At the start of the test condition, just after the pilot disengaged the autopilot, the lower rudder deflection diminished about 10% to -4.5 degrees, as the aircraft began a yaw and rolling motion to the right. Sideslip remained roughly constant, though fluctuating, during this brief interval as the lower rudder deflection diminished. The lower yaw damper may have commanded the 10% decrease in rudder deflection in response to the onset of right yaw. In contrast, the upper rudder deflection remained at -3.5 degrees. Because the plots, provided by the manufacturer, did not include a time history of rudder pedal force, the analyst can not be certain of the exact inputs made by the pilot. However, from the traces of rudder surface position it would appear that yaw damper was active at FTI time 13:43:10 through 13:43:19. These apparent yaw damper inputs consisted of a 2 degree right rudder deflection, followed two seconds later by a 0.5 degree left deflection, followed 1.5 seconds later by a 1.5 degree right rudder deflection. The manufacturer did not include instantaneous yaw rate on the data plot. Examination of the INS heading trace shows that these rudder (yaw damper) deflections occurred at slope changes of the heading trace.

The test instrumentation for the upper rudder may not have been accurate. A comparison of the upper and lower rudder traces after time 13:43:25 suggests that the FTI recorded neutral position for the upper rudder as about 0.3 degrees left deflection. Thus, an analyst should make a correction of -0.3 degrees to the indicated values for upper rudder.

The sideslip trace, plotted as differential pressure in units of inches of water, varied from 6.5_ during the sustained left sideslip, to 1.5_ twenty seconds after the rudder release. Following the test condition, after the release of the rudder pedal, there are several peaks and dips shown on the sideslip trace. The analyst can compare these local maxima and minima with those shown on the rudder position traces. The slope reversals on sideslip trace lag those of rudder (yaw damper) trace by 1 to 2 seconds. Had the NTSB ordered a more thorough “data request,” the time history of sideslip angle might have been a more useful parameter.

The Board included the flight test data discussed above in the NTSB docket. The data was part of the Addendum to Performance Group Report, dated January 7, 1981.

Proposed new subSection 1.16.5, TWA TESTS ABOARD N840TW . The TWA flight tests of 1977 included an incident that involved aircraft controllability. The incident occurred after a suspected autopilot false disconnect. The pilots encountered heavy stick forces, uncommanded control inputs, and the uncentering (ratcheting) of the aileron control wheel. The Board did not previously consider this anomaly of the B727 flight control system. It constitutes new evidence. The Board should include in the docket the enclosed affidavit from the eye witness of that 1977 incident. The Board should also require that airlines forward to the NTSB details of past B727 control incidents, or incidents of high altitude instability. A proposed new sub section, 2.5.3 (titled Autopilot Roll Channel) includes further analysis relating the 1977 TWA test results to the accident of TWA 841. The enclosed affidavit describes that 1977 incident.

The affidavit describes a flight controls incident that had occurred previously aboard the accident aircraft.

_

In Re TWA Flight 841,

April 4, 1979,

Near Saginaw, Michigan

AFFIDAVIT

During a crew debrief of the flight controls incident that occurred on 17Dec89 aboard TWA Flight 1070, I, Capt. P. T Williams (TWA), was prompted to recall a problem with the B727 flight control systems that had occurred years earlier.

*********************

I, Capt. P. T . Williams (TWA), witnessed an incident that occurred during a flight aboard the B727-100, plane N840TW, on Flight 5562, from MCI, on May 23, 1977. At the time, I was a TWA Check Airman. I was flying the aircraft from the left seat (but not as PIC) while receiving a Proficiency Check Ride from Capt Sal Fallucco (who was Pilot-in-Command). Oddly, also aboard that day, as observers, were several TWA employees involved in engineering, and pilot training. During the flight I was directed to forcefully overpower the roll channel of the engaged Autopilot, so that the observers could evaluate the response of the aircraft - autopilot systems. As a result of such repeated test conditions, an abnormal flight control incident developed.

I was directed to disconnect the autopilot during these “overpower” conditions. During the first test condition the autopilot cleanly disconnected. The difficulties began when during the second test condition, the autopilot apparently disconnected (the A/P Paddles dropped to the Disengaged Position) but the flight controls became extremely difficult to manipulate in roll, and there was no annunciation provided to indicate the source of the problem. After much effort on the part of the pilots the aircraft was returned to MCI. With emergency equipment standing by, a precautionary landing (at flaps 5 degrees) was made by Capt. Wayne Disch -- Manager of Training at that time. During the landing rollout at MCI, when aileron inputs were no longer needed, Capt. Disch released the control wheel, which then rotated full travel. Something then snapped beneath the floor of the cockpit, and the control wheel returned suddenly to a normal position. Thereafter the control wheel was free, and could be rotated normally in either direction.

*******************

The foregoing is true to the best of my knowledge and belief.

[Signed]

P. T. Williams

Captain

Trans World Airlines

Subscribed and sworn to before me this24thday ofMay, 1990.

[Signed]

Michelle Dunnavant

Notary Public--State of Missouri

Section 1.17.1 -- Under B-727 FLAP SYSTEM , on page 14 of the NTSB AAR, the first sentence was erroneous. It incorrectly stated that there are four LE Flaps. In fact, there are six LE Flaps, three on each wing of a B727.

Section 1.17.2 -- HISTORY OF B727 LEADING EDGE SLAT PROBLEMS ,NTSB AAR page 18. The NTSB failed to mention that B727 aircraft had suffered numerous cases of slat separation. Prior to 1979 SDR’s showed a total of sixteen documented cases of loss of a slat surface. The documented cases indicated that the crew usually noted the loss of “A” Hydraulic System immediately after the loss of a slat. (In eight other cases poor documentation proved insufficient to determine whether the extended slat had actually separated.) These cases involved uncommanded extension of the slat prior to separation. Data included in the SDR’s and incident reports verified the manufacturer’s calculations regarding the limited load bearing strength of an extended slat. We discuss this subject in a proposed analysis section (2.4.3),titled “Limited Strength of an Extended Slat.”

Section 1.17.3 -- Under AIRCRAFT PERFORMANCE ,on pages 18 through 20 of the NTSB AAR, the Board based their analysis on the false assumption that a LE Slat extended at FL390. Paragraphs that discussed the calculated rolling moment, due to an extended slat, were not pertinent. A reader could mistake these calculations for fact. The Board should delete these paragraphs from the NTSB report (or at least remove the paragraphs from the factual section of the report, and insert them into the analysis section). The last paragraph on AAR page 19, and the next three paragraphs on page 20, discussed such calculations and are not pertinent to the upset sustained by the accident aircraft.

A graph was displayed on AAR Page 21 that depicted the rolling moment due to an extended #7 slat. The graph is not pertinent to the upset of the accident aircraft and should be deleted from the NTSB Report.

Section 1.17.4 -- titled NO. 7 LEADING EDGE SLAT OPERATION , was on pages 20 through 22 of the NTSB AAR. The two existing paragraphs that form this subsection of the NTSB report described the loads on a retracted slat under the initial upset conditions (FL390 cruise). The discussion was incomplete because it failed to mention the factors that would interact to pull a slat from the retracted position following loss of Hydraulic System “A” pressure. We present these factors in a proposed new (sub) Section 1.17.5 titled “Loads On A Slat.”

The Board should use this Section 1.17.4 to better describe certain aspects of the operation of the leading edge slat.

On page 16 of the NTSB AAR, there was an illustration that showed a side-view of a slat. The illustration showed the slat in the retracted position, and in a mid-extension in-transit position. Unfortunately that illustration was not thorough enough. The illustrator failed to label several components important to the proper operation of a slat.

Engineers presented a better drawing in the manufacturer’s report on page B-19. The Board should include this drawing in the NTSB report and refer the reader to it in Section 1.17.4. This drawing has a more complete depiction of the slat components, and a more thorough identification of those components. This drawing offers the reader a view of an extended slat and the interaction of important components.

The Board should describe the function of theSLAT ACTUATOR RETRACT LOCK INDICATING SWITCH. The Board must tell the reader that a fault in this indicating switch could leave the pilots with no means of identifying an unlocked slat, following slat retraction after takeoff. Normal hydraulic pressure would hold an unlocked slat actuator piston in the retracted position. However, any subsequent loss of hydraulic pressure could result in the slat being pulled from the retracted position by various forces acting on that wing section.

The inboard track of the # 7 slat suffered a structural failure of itsEXTEND STOP.[13]The Board should describe and illustrate the function of the components of that slat track and the slat track carriage. The Board should describe the impact overload forces encountered by the inboardSLAT TRACK EXTEND STOPwhen it contacted theTRACK CARRIAGE STRUCTURAL STOP.

The Board should describe the actuation of theSLAT EXTEND POSITION INDICATING SWITCHas the slat track moved to the extended position. The reader should understand that even after the #7 slat separated from the aircraft, the annunciation failed to alert pilots that the #7 Slat was anyway abnormal. Lacking any input from the faultedSLAT ACTUATOR RETRACT LOCK INDICATING SWITCH, the amber “LE Flap” annunciation would fail to illuminate when the #7 slat unlocked. The green “LE Flap” annunciation would have indicated that #7 slat extended after repositioning the flap handle during the Alternate Flap Extension Checklist.

Proposed new subsectionSection 1.17.5, LOADS ON A SLAT .This revised subsection should discuss both tensile and compressive loads on the actuator of a retracted slat, and the load limit of an extended slat. The Board should describe how such factors as high speed, high G’s, vibration, and sideslip, influence loads on the actuator of a retracted slat. Then the Board could describe the effects of those factors on the structural components of a slat subjected to sudden over extension following loss of hydraulic pressure.

Loads on an Extended Slat

The Board should discuss the design load limit of an extended slat. Consider this hypothetical question: Had the #7 slat extended at FL390, at what point during the uncontrolled dive would it have failed and ripped away from the wing? Using design coefficients, substantiated by flight loads survey data, the manufacturer calculated the “most probable” moment of slat separation to occur at a drag load of 5.11 psi. (The manufacturer presented that data as a graph on page B-12 of their report to the NTSB dated Sept. 24, 1979.) The manufacturer determined that the accident aircraft encountered this most probable failure load at 363 KIAS. By referring to that point in the FDR tabular data, a rate-time interpolation between altitude data points shows that the accident aircraft exceeded that airspeed while diving through 31812 feet FDR altitude. (After correcting for the known FDR altitude error, the slat could have stayed intact until an actual altitude of about 31500 feet.) The manufacturer stated, on Page B-5 of their report, that the probable time that an extended #7 slat would have separated due to overload was 24 minutes, 1.5 seconds (Flight Data Recorder time). Had an extended slat caused the upset and the uncontrollable rolling moment, the slat would have separated during the first third of the dive.

The aircraft would then have been controllable. Yet the accident aircraft continued to dive uncontrollably for another thirty seconds. An early separation of the slat, along with the associated fracture of the slat actuator, would have resulted in loss of pressure in the “A” Hydraulic System. Yet that hydraulic system was intact nearly thirty seconds later in the uncontrolled dive, when it powered the landing gear extension. The trail of debris found on the surface suggested that the #7 slat could not have separated until after gear extension. Correlation of the evidence suggests that an extended #7 slat could not have caused the sustained interval of uncontrollable flight.

The Board should carefully consider one further point. When estimating the most probable moment of failure and separation of an extended #7 slat, the above estimate of 363 KIAS is probably unrealistic because of the worn slat components involved in this case. The manufacturer arrived at that estimate by comparing air loads with slat design loads. The manufacturer’s report (on page B-4) alludes to this inadequacy,

Intact slats would have survived the initial seconds of the incident in an extended configuration. It is apparent from physical evidence, however, that slat #7 was not intact . . . due to pre-existing damage.

The manufacturer’s own metallurgist documented slat misalignment and fatigue progression in components. Had the #7 slat actually been in the extended position from the very beginning of the upset, the “most probable” moment of separation would occur prior to the instant calculated above. The worn slat would have ripped away even earlier within that first third of the uncontrollable dive segment.

LOADS ON A RETRACTED SLAT

The manufacturer’s report also provided the NTSB with a suggestion of how and when that #7 slat ripped away from the accident aircraft. The following excerpt (manufacturer’s report, Page B-2) addressed the conditions under which one or more slats could have extended;

(a) if the slats were unlocked, hydraulic system off, spoilers up and/or load factor less than 1.0; or if,

(b) inadvertent or intentional partial extension occurs under conditions of high mach number and low lift coefficient.

Each of the variables identified in the above excerpt contributed to thelocalized forces acting on a retracted #7 slat. Yet in this case, perhaps other variables contributed to the local aerodynamic forces. Various conditions affected the outboard section of the right wing. The aircraft was in a left sideslip (nose-right, left wing forward), which induced proverse roll to the right. [14]During most of the dive, the aircraft was experiencing high speed buffet. While diving through 460 KIAS, M.85, about 12,000 feet, the gear extension caused damage to hydraulic lines. Each slat actuator lost the hydraulic pressure that had forced its piston to the retract position. Air loads on the right outboard aileron forced it to float from its normally faired position. This up-float, over one inch upward, affected that retracted #7 slat as would a deflected spoiler. Flutter of the right outboard aileron may have excited the wing section into vibration. The evidence suggests that these conditions existed for a very short time during the sequence of failures. These conditions existed immediately after the right main landing gear suffered damage as it over extended due to the side loads. And the evidence suggests that it was at this point in the failure sequence that the #7 slat extended as the hydraulic pressure to the retract side of the slat actuator decayed.

Proposed new sub section ,Section 1.17.6 ,B727RUDDERCONTROLSYSTEMS.

The following information about the rudder control systems is provided to the pilots in the TWA Flight Handbook.

The upper rudder is powered by [hydraulic] system B and the lower rudder is powered by [hydraulic] system A. There is no manual operation of the rudders; however, the lower rudder can be powered by the standby hydraulic system through a separate actuator. . . . A cable system transmits rudder inputs from either pilot’s rudder pedals to the power units in the vertical stabilizer. [15]

The yaw damper system is designed to counteract yaw due to dutch roll. Independent systems on the upper and lower rudder sense yaw and position the rudders to stop it. The yaw dampers are designed to operate full time. [16]

At high altitudes and cruise mach numbers, the Dutch Roll characteristics of the 727, without the yaw dampers, is undamped and divergent. If not corrected it will deteriorate into a complete loss of control. [17]

Lacking the stability augmentation provided by the yaw dampers, the high altitude lateral dynamic instability characteristic of the longer B727-200 aircraft isNOTequal to the dynamic instability characteristic of the shorter B727-100 aircraft. The Dutch Roll characteristic of the shorter model is less damped. The fraction of critical damping decreases to zero as the altitude increases toward FL260. Above about FL260, for the B727-100, the fraction of critical damping changes sign and becomes more negative (divergent) at higher altitudes. (This data was disclosed in attachment D, of a letter from H. P. Hogue of Boeing, to R. VonHusen of the NTSB; dated December 19, 1980.)

The normal Power Control Unit (PCU) for each rudder serves acommon mode function . Each hydraulic powered rudder is positioned by a PCU which responds to commands from the pilot through his rudder pedals. Concurrently, each PCU is designed to respond to electrical signals input from its associated Yaw Damper (a subsystem composed of numerous parts, located in various sections of the aircraft, and accepting inputs from several other subsystems). The two input commands to each rudder PCU -- the rudder pedals and the yaw damper -- are completely independent. There is no feedback to the pilot’s pedals when a rudder is displaced by a yaw damper command.

The flight test of October 2, 1980, demonstrated the maximum deflection angle of the rudders for the cruise conditions at FL390. Full rudder pedal travel yielded a rudder surface deflection angle of: ru = -3.5° for the upper rudder; and rL= -4.5° to -5° for the lower rudder. Instrumentation recorded those measurements after the aircraft was established in a steady nose right sideslip using full right rudder pedal travel. (The Board failed to determine the exact value of the sideslip angle during that important test condition.)

For this case of nose right sideslip, hinge moment calculations show much less rudder surface deflection available in the opposite direction, acting to oppose the sideslip. The negative sideslip angle, airplane nose right, acts to limit the magnitude of the available positive (left) deflection of a rudder surface. The rudder hinge moments increase proportionally with the square of the equivalent airspeed, thus also limiting the available rudder deflection

A malfunction of either or both rudder control systems could result in split rudders. The manufacturer (in their Document D6-8095 pages 4.3-1 through 4.3-6) considered such rudder anomalies:

Due to the several boost systems and the dual pressure available on the main systems, there is a multiplicity of failure modes. . . . It should be noted that there is a considerable difference in rudder hinge moments between normal operation (both segments deflected) and abnormal operation (segments deflected separately). [18]

Utilizing data in the above reference, investigators could calculate possible abnormal rudder deflection angles. Consider a case of a nose right sideslip: with   at 230 KEAS, at FL390; with a malfunctioning lower rudder actuator inputting maximum right deflection. At normal hydraulic boost pressures the lower rudder could achieve approximately rL = -4.28°deflection (right). With full left rudder pedal input, opposing the negative sideslip angle, the hydraulic actuator could only deflect the upper rudder surface to approximately ru = +2.12°(trailing edge left).

It is difficult to model the rolling moments and yawing moments resulting from possible aberrant rudder deflections. Calculated control margins may ignore the additional moment increments that ensue from dynamic effects of roll rate and yaw rate. The manufacturer attempted some simulator testing following the accident. However, their simulator emulated the stretch model of the aircraft — a 727-200. The TWA 841 accident involved a 727-100 aircraft which has significantly different directional and lateral stability characteristics (especially at high cruise altitudes). The manufacturer consistently reassured the Board that the characteristics of the two models were similar. However, implicit in several communications from the manufacturer was the suggestion of critical differences. In a letter dated 19Dec80, from Boeing’s H.P. Hogue to the NTSB’s R. VonHusen, the manufacturer conceded that,

The 727-100 would have more sideslip due to rudder than would the 727-200 . . .

The manufacturer explicitly associated that information with its effects on the control margin of an aircraft with an extended #7 slat. The Board failed to recognize the real importance of that sideslip evidence. Sideslip and the rudder control systems were important variables, critical to the high altitude characteristics of the B727-100.

The manufacturer worked with the NTSB’s investigators and they consistently judged new evidence with such a bias. The investigators had constructed an interpretation of an ambiguous situation. Thereafter they routinely processed new information with a bias toward their interpretation. The NTSB’s investigators either adopted or abandoned evidence, based upon that bias. [19]The investigators’ erroneous assumption, that an extended slat had caused the upset, prevented them from properly documenting more pertinent evidence.

The controllability testing and related calculations accumulated during the NTSB’s investigation applied to an aircraft flying at FL390 with an extended slat. The NTSB’s records do not include sufficient information about faults in the lateral and directional control systems that may have contributed to spiral divergence. Lacking information, it is difficult for any analyst to verify the possible control margins during a fault in one or both of the rudder control systems. (The Board should study the investigations of the AA Flight One accident, and the MAC 59402 mishap, to learn appropriate investigative methods for such control system failures.)

The B727 Maintenance Manual (MM 22-00, Page 1) stated that yaw damper commands have an authority limit of ±5 degrees of rudder deflection.

The B727 Maintenance Manual provided the investigators with more information about the yaw damper commands, and the rudder system:

The yaw damper consists of two yaw damper couplers, two guarded yaw damper engage switches, a rudder trim and position indicator and two rudder position sensors. . . . Each of the two yaw damper couplers sends electrical signals to their respective rudder power control packages. . . . When the yaw damper is engaged, the yaw rate gyro senses any changes in the yaw axis, and the yaw damper provides the necessary airspeed compensated signals to the rudders to stabilize the airplane. [20]

Two yaw damper couplers are installed in the electronic equipment racks, one for the lower and the second for the upper rudder. Each coupler consists of a yaw rack assembly and several plug in modules. The plug in modules consist of the yaw rate gyro, yaw servo amplifier, yaw synchronizer, yaw calibrator . . . .

Each rate gyro is a non-hermetically sealed unit with simplified in-line construction. No slip rings or brushes are used for the microsyn stator, microsyn rotor, or the hysteresis motor stator. The gimbal assembly is supported by permanently lubricated ball bearings and a torsion bar whose movement is limited by a stop mechanism. A viscous damper is used to give an adequate damping ratio to the gimbal assembly.

The electronic plug in modules contain several electronic cards. . . . Each card contains the electronic or electro-mechanical components needed for amplification, synchronizing, or band pass filtering of the yaw damper signals. Each yaw rack assembly contains a 30 volt dc power supply and a transformer providing the various ac excitation voltages. The yaw calibrator contains various resistors to adjust the gains of the yaw damper coupler to a specific airplane configuration. [21]

The sensors for each yaw damper consist of the yaw rate gyro and the linear series actuator position transducer. The yaw dampers are engaged by separate Rudder Yaw Damper Engage switches. The yaw damper electronics are contained in the yaw damper coupler which comprises the following functional modules for signal coupling, shaping and amplification: yaw synchronizer, yaw servo-amplifier and gain calibrator. The air data sensor varies yaw damper gain as a function of airspeed (dynamic pressure Q-parameter control). The rudder servo system is composeed of a rudder power unit which contains the transfer valve, the yaw damper actuator, the actuator position transducer, the pilot input linkages and main control valve and actuator. [22]

The signal input circuitry, transducer position feedback, wipe-out circuit, filters, and signal summing circuits, are described in the MM 22-00, on page 19.

The Rudder and Elevator Position Indicator as installed on the center instrument panel of TWA B727 aircraft is different than that used on aircraft at some other airlines. There is a position pointer for each rudder, intended to display rudder deflection. The indicator also has a failure alert flag for each yaw damper. The materials provided to flight crews describe this fail flag as an annunciation for loss of ELECTRICAL power to the yaw damper, or the engage switch toggledOFF. As installed on TWA aircraft this failure flag signals loss of electrical power only. The flag does not annunciate the loss of hydraulic pressure to the rudder actuator. Such a failure alert is only displayed for the crew, and not recorded by the flight data recorder.

The lower Yaw Damper Fail Flag was observed by the flight crew of accident aircraft. They first noticed this failure alert after recovery from the uncontrollable flight segment. The NTSB never determined why the fail flag appeared. Investigators may have mistakenly assumed that the Yaw Damper Fail Flag was a consequence of the loss of hydraulic pressure to the lower rudder.

Section 1.18 -- titledUseful or Effective Investigative Techniques , on page 22 of the NTSB AAR. The Board adopted a deceptive investigative technique suggested by the manufacturer. The fourth sentence erroneously states,

This technique permitted the illustration of highly accurate g-trace frequencies . . . .

The data recorder mechanisms each had a limited frequency response. The data recorder mechanisms record all vibration of any frequency greater than about 6 cps at that limit for frequency response. This was an unprecedented and an unrecognized forensic technique. Furthermore, such a technique of vibration frequency comparison was inappropriate in this case due to the limited frequency response of the particular recorder mechanisms. The results from such a vibration comparison were therefore ambiguous. Yet the Safety Board applied this comparison technique in an attempt to correlate data from an accident with data from one prearranged test condition. By using this inappropriate comparison technique, investigators were able to find apparent vibration similarities between unrelated conditions.

Proposed new sub section 1.19,Efforts to Revise Investigative Errors . Newspaper articles and a documentary on network television -- about the mistakes in the NTSB’s investigation of the accident -- prompted an independent citizen to suggest an alternative hypothesis for the upset. In a paper offered to the Accident Investigation Department of the Air Line Pilots Association, Mr. Duane Yorke, of Massapequa, N.Y. described the results of his own analysis.

The strength of Mr. Yorke’s analysis was that it made sense of the evidence. Damage to various components of the accident aircraft fit with the observations offered by witnesses (the pilots) to suggest a failure sequence. Subsystem failures initiated the upset and sustained the uncontrollable flight segment. Then a final subsystem failure resulted in controllable flight. His inspiration for this analysis was a knowledge of previous yaw, roll-over, vertical dive incidents during the early days of swept wing aircraft.

Mr. Yorke has thirty years of experience in aeronautical engineering. He holds a BSAE from MIT. He served as Director of Supersonic Aircraft Development for Grumman Aircraft Corporation.

The Board should rearrange the Analysis portion of their report, so that they can present the analysis of the direct evidence in a logical manor. The analysis should begin with the trail of debris found on the ground north of Saginaw. The report should then “work backwards” from the recovery phase to the initial upset, correlating the direct evidence. The Board’s Report should correlate the specific damage sustained by the accident aircraft, with the information from the Flight Data Recorder, and the observations of the crew and passengers.

The direct evidence fits together to suggest a sequence of failures. The accident analysis should focus on direct evidence. The Board should carefully avoid the use of assumptions and circumstantial evidence. The Board should correct those errors in its analysis in which they ignored direct evidence. When the manufactured circumstantial evidence contradicted the direct evidence, the Board often employed the circumstantial evidence and rejected direct evidence. For example, the Board consistently rejected testimony of the crew in favor of conflicting circumstantial evidence.

The erroneous assumption that the aircraft experienced a slat extension while at FL390 tainted major portions of the Board’s Analysis Section. In order to correct the resulting mistakes in their report, the Board should now discard many whole paragraphs from their initial report.

Section 2.3 --The Aircraft , on pages 22 and 23 of the NTSB AAR. Two paragraphs described the accident aircraft as having been free of maintenance discrepancies, and stated that the flight crew had not noticed any malfunctions during the takeoff and climb-out. However, the Maintenance Records Group identified two discrepancies found during the previous “C” Check, which may have persisted and contributed to the later accident. The #7 slat actuator and the lower rudder actuator both may have been defective. During the previous “C” Check, maintenance inspectors documented hydraulic leakage near each of those actuators. Evidence suggests that each of those units experienced faults during the accident that occurred one month after the “C” Check.

The Board should add an additional paragraph discussing the lack of fault annunciation for many (normally passive) subsystems on that B727-100 model of aircraft. Most subsystems are poorly instrumented. In the design of such units the manufacturer provided no direct method to alert the pilots that a subsystem may be unreliable. Some examples of subsystems that lack fault annunciation in the cockpit are: the cockpit voice recorder; either yaw damper; either rudder actuator; a slat actuator retract locking mechanism and indicating switch. Thus, without any form of fault annunciation, a flight crew could only identify a fault in a subsystem after experiencing an “incident.” During such an event at least one crew member must first perceive that the faulted (normally passive) subsystem had failed to perform an intended function. Only after such a subtle discovery could a crew member attempt to identify the particular fault .

Proposed new analysis section 2.4,DIRECT EVIDENCE .

In the Board’s initial Report, Section 2.4 was titled “Extension of the No. 7 Leading Edge Slat.” That section had been a major part of their report because the Board regarded the slat extension as part of the initial upset. That section of their report included eighteen paragraphs, covering various topics, spread over pages 23 through 27 of the NTSB Report (AAR).

The Board should completely rearrange Section 2.4 of their report. The slat separation was only an effect of the upset, not a cause. The Board could utilized the Section 2.4 of their revised report to present analysis of the direct evidence.

There were numerous errors in the NTSB’s investigation. Some of the analysis was completely inappropriate, done in an attempt to support the Board’s erroneous assumption that a slat extended at FL390. Many of the paragraphs included in that Section 2.4 of the Board’s initial Report are discussed in the proposed new sections 2.6 and 2.7 and 2.8.

Proposed new sub section2.4.1,THE TRAIL OF DEBRIS .

The wreckage distribution chart is one of the most useful tools the investigator can use . . . failure patterns and failure sequences suggest themselves when the completed distribution chart is carefully studied. . . . the wreckage distribution serves as the only record of how the various pieces were located at the accident scene. The significance of later findings often depends upon reference to the original wreckage distribution chart; and if one had not been prepared, the investigation could be seriously hampered. [23]

Several aircraft parts ripped away from the aircraft in-flight. These aircraft parts were found impacted on the surface north of Saginaw. Included in the manufacturer’s report on page A-10 was Figure 4. (See the illustration reprinted on the next page). This figure showed that four parts ripped from the aircraft had impacted the ground along a path stretching five thousand feet, oriented to a northeast to southwest baseline. Most importantly, this figure revealed something not precisely stated in the final NTSB report: the separated flap track “canoe” fairing[24]impacted very near the inboard half of the #7 slat. This Figure 4 showed that the outboard half of the #7 slat impacted about one thousand feet from the previously mentioned parts, as measured along the reference baseline. About four thousand feet further to the southwest, a portion of the #6 flight spoiler panel6 flight spoiler panelimpacted.

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Trajectory Analysis

of Separated Parts

The flap track “canoe” fairing referred to above was discussed later in the structures section of the report from the manufacturer (Page B-17):

. . . the flap track fairing that came off the right hand inboard flap. . . . There was nothing evident that the departure was caused by other than abnormal loads occurring in the maneuvers of the airplane, probably when the gear was extended.

The inboard half of the #7 slat impacted near that flap track “canoe” fairing. Further analysis of the manufacturer’s trajectory chart is proposed in a later section (2.4.7). This analysis of the trajectory of the separated parts suggests that the inboard half of the slat also ripped away from the aircraft as the right gear extended. [See the graphic overlay of the aircraft profile track and the parts’ trajectory, page 63 herein.]

Proposed new sub section2.4.2,GEAR EXTENSION DAMAGE AND HYDRAULIC SYSTEM FAILURE

At some point during the uncontrolled dive the “A” System hydraulic pressure decayed. The crew identified this hydraulic system failure after the recovery.

During the uncontrolled rolling-dive the crew extended the landing gear using the normal means of extension. Hydraulic System “A” pressure successfully removed the gear up-locks and powered the normal gear extension. (Thus, the #7 slat could not have separated until after gear extension. A fractured slat actuator would have resulted in loss of pressure in the “A” Hydraulic System, which would have disabled the normal means of gear extension.)

During the post-flight inspection investigators found damage to the System “A” hydraulic line near the right main landing gear. Damage resulted when the right main gear over-extended.

The “A” Hydraulic System was intact and pressurized throughout the uncontrollable upset and dive. The “A” Hydraulic System failed as the right main gear swung to an over-extended position. The extreme force of this extension swing caused structural failures to the right main gear side brace and actuator support beam. Furthermore, this structural damage ruptured the adjacent hydraulic fluid cooling line. This rupture, in a line pressurized to 3000 psi, caused a sudden failure of the “A” Hydraulic System.

The crew reported that the aircraft was not controllable during the upset and rolling-dive until the landing gear extended. Then the controls again were effective.

The flight controls became effective again as a direct result of the loss of “A” System Hydraulic pressure through the ruptured line, near the right landing gear. This suggests that one of the actuators, powered by the “A” System Hydraulics, had malfunctioned and caused an unexpected displacement of a flight control—the lower rudder. Loss of hydraulic pressure to that malfunctioning actuator then permitted the flight control to center.

The fortuitous sequence of failures that restored controllability are illustrated on the next five pages. These illustrations also relate the trail of debris to the failure sequence.

__

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Landing Gear Extends

Extends

(sideslip)

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(landing gear side brace, landing gear damage, hydraulic leakage, flap track “canoe” fairing hydraulic line rupture

(landing gear side brace, landing gear damage, hydraulic leakage, flap track “canoe” fairing hydraulic line rupture

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Recovery--Failure Sequence

hydraulic leakage, lower rudder, flap track “canoe” fairing, slat #7 hydraulic line rupture )

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Continued Failure Sequence

(landing gear damage, spoiler damage, speed brakes, spoiler blow down, separated parts slat #7access panel damage)

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The Recovery and Trail of Debris

( landing gear damage, slat #7, hydraulic line rupture, rudder, moon, separated parts, trail of debris )

Proposed new sub section2.4.3,LIMITED STRENGTH OF AN EXTENDED SLAT .

This portion of the analysis should discuss “Factual Information” proposed for inclusion in the revised NTSB Report under a new subsection 1.17.5, “Loads On A Slat.”

The Board assumed that the #7 slat extended while the aircraft was cruising at FL390. That assumption resulted in the Board’s erroneous conclusions. Building on that erroneous assumption, the Board further assumed that the extended #7 slat initiated the upset and caused the sustained interval of uncontrollability.

The Board stated an erroneous conclusion near the end of the “Analysis” part of the NTSB Report, at the top of page 33.

The aircraft continued to descend rapidly, and it continued to roll to the right until the #7 slat was torn from the wing and lateral control was restored. . . . at an altitude of about 8000 feet.

The Board’s assertion regarding an extended slat contradicted data presented by experts employed by the manufacturer. That data described the structural strength of an extended slat. As presented earlier in the proposed section 1.17.5, the manufacturer’s experts calculated that an extended slat could not have sustained loads beyond the first third of the uncontrollable flight segment. An extended slat would have ripped away as the accident aircraft accelerated through 363 KIAS during the dive.

Furthermore, as presented as “Factual Information” in the proposed section 1.17.5, the metallurgists found pre-existing misalignment and fatigue progression in the slat components. Had the #7 slat actually been in the extended position from the very beginning of the upset, the “most probable” moment of separation would occur prior to the instant calculated above. The worn slat would have ripped away even earlier within that first third of the uncontrollable dive segment, at a higher altitude and a slower airspeed than that calculated by the manufacturer. However, the wreckage distribution suggests that that slat did not separate until much later in the dive, after the landing gear extended.

Failure data from other in-flight incidents corroborated calculations that an extended #7 slat is likely to rip-away at speeds near 363 KIAS. The extension and separation of the #2 or the #7 slat had occurred during numerous in-flight incidents prior to the 1979 TWA 841 accident. The separation of a slat, immediately following an uncommanded extension, had occurred at speeds of 250 KIAS or less (EAL/N8120N on 13Feb73). Following another uncommanded extension, at 235 KIAS, the slat box buckled at mid span and the slat tracks bent aft (EAL/N8133 on 24Mar70, this slat remained attached). In seven of the other incidents, [25]a #2 or #7 slat extended and separated at speeds between 310 KIAS and 360 KIAS. Prior to 1979, the SDR’s showed a total of sixteen cases that involved loss of a slat surface.

For the case of the accident aircraft, we analyze failures of the structural components of the #7 slat later in the proposed section 2.4.6. Analysis shows that the slat actuator rod fractured immediately upon the extension of the slat. Therefore the slat actuator rod did not contribute to the load bearing strength of the extended #7 slat.

The Board should acknowledge the following deduction: During most of the upset and dive, hydraulic pressure held the #7 slat in the more secure, retracted position. The #7 slat remained in the retracted position until a rupture occurred in hydraulic lines near the right gear strut, as the right main gear swung past the normal position. We analyze the fault in the slat Retract Lock and Retract Lock Indicating Switch later in the proposed sections 2.6 and 2.7.

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Insert -- Illustration, Limited Strength of an Extended Slat

( Slat load bearing strength, FDR speed verses slat strength, )

Proposed new sub section2.4.4,EVIDENCE OF SUSTAINED SIDESLIP

The aircraft sustained differential damage during extension of Main Landing Gear. The right main landing gear sustained extensive damage; the gear side brace broke during the over extension, as did the actuator support beam. The right gear strut swung past a normal “extended” position causing impact damage to the most inboard flap track. The flap track “canoe” fairing on the inboard track of the right inboard flap separated from the aircraft. The left gear structure was NOT similarly damaged and was essentially unharmed.

This pattern of differential damage suggests that at the time of gear extension, a large left wing forward, nose right sideslip angle existed. As each main gear extended, it swung outboard. The nose right sideslip, with the relative wind coming in from the left, resulted in the severe side load on the right gear as it swung down.

The existence of this apparently substantial sideslip angle at the moment of gear extension, at such high speed and high ‘G’ conditions, requires a careful consideration of the possible causes of the sideslip condition.

Some crew members who have experienced sideslip conditions have noted the shift of the noise level in the cockpit. Most B727 pilots are familiar with the short intervals of increased noise level in the cockpit that occur during penetration of moderate turbulence. These changes in cockpit noise level result from brief sideslip episodes. In cases of more prolonged episodes with greater sideslip angle the variation in cockpit noise level is more pronounced. For example, in the similar mishap of MAC 59402,[26]at least one pilot remarked that during the sudden upset at 41000 feet the noise level in the cockpit was deafening.

The flight engineer, aboard TWA 841, also described a sudden change in sound, noted during the upset recovery:

Nothing was working . . . And I thought we were going in. . . . But I remember that when the gear went down, I really hadn’t heard anything that loud explode in the plane . . . the control surfaces were now taking effect. . . . You could hear a complete change of the plane. [27]

This change of noise level resulted from the cessation of the sideslip condition that had persisted prior to the gear extension.

Proposed newsubsection2.4.5,PRE-EXISTINGLEAKAGEFROMA WORN ACTUATORFOR#7SLAT

Maintenance records for N840TW showed that the airline had accomplished the most recent “C” Check about one month prior to the accident. At that time inspectors noted a “very heavy accumulation of skydrol” that had leaked into the wing leading edge, aft of the slats. The mechanics recorded their “corrective action” for that discrepancy as: “tightened . . . the ‘B’ nuts outboard of the #6 and #7 slat actuator[s].” The mechanics then washed the area and determined there was no leakage from those “B” nuts. [28]

In addition maintenance records from the most recent “C” Check showed that the skydrol had damaged (stained) the silver finish on the upper surface of the right wing. During the “C” Check the mechanics used paint to restore the silver finish on top of the right wing. [29]

After the accident, during an inspection of the right wing upper surface, investigators recorded evidence of skydrol bathing (staining) of an area near and behind the No. 7 slat position. [30]

The Board should conclude that there was persistent leakage of skydrol that continued even after the “C” Check. During the “C” Check the mechanics had tightened the plumbing connections and verified that those connections did not leak. Yet, skydrol leakage continued and over a period of several weeks again damaged the silver finish on the upper surface of the right wing aft of the #7 slat. The Board should conclude that the persistent leakage originated in the #7 slat actuator itself.

Boeing Operations Manual Bulletin 75-7 stated that external leakage of the slat actuator was one of the clues used in identifying actuators with damaged retract lock rings. Several other anomalies inside the actuator retract locking mechanism could also have resulted in such leakage. We discuss this further in the proposed analysis Section 2.6, “Fault Analysis of Slat Retract Locking Mechanism.”

Proposed new sub section2.4.6,EXTENSION OF THE NO. 7 LEADING EDGE SLAT

For documentation of pre-existing damage and misalignment of the #7 slat refer to Section B of the manufacturer’s report and page 33 of NTSB AAR.

A variety of forces acted on the components of the retract locking mechanism for the #7 slat actuator during the segment of uncontrollable flight. The combined effects of the following forces acted on the retract lock: vibration, unusual G-loads, hydraulic pressure surges, and loss of hydraulic pressure. Furthermore, aerodynamic pressures across the surface of the slat acted through the actuator rod, creating unusual dynamic forces on the retracted piston inside the head end of the slat actuator.

Boeing Operations Manual Bulletin 75-7 advised crews about factors which might result in the uncommanded extension of a slat. The problem could occur with speed brakes deployed at speeds above M.80. The loss of hydraulic pressure to the slat actuator could permit various forces to pull the slat from the retracted position, if the retract lock mechanism failed.

During the last moments of the uncontrolled flight segment, a dynamic and unusual localized aerodynamic pressure distribution acted across the surface of the retracted #7 slat. The up-float of the right outboard aileron and the nose right sideslip angle of the aircraft contributed to the unusual pressures over the surface of the slat.

The Board should conclude that the various tensile forces resulting from the conditions outlined above pulled the #7 slat from the retracted position. This extension of the slat occurred at approximately 24:31 FDR Time, immediately following the damage sustained as the landing gear extended. The slat tracks reached the extension limit with considerable energy. During the post-flight inspection, investigators found the inboard track Extend Stops sheared-off and lying inside the enclosed wing section.

The now extended #7 slatwas probably the weakest of the slats, as evidenced by the misalignment and the defective inboard T-bolt connection. The #7 slat promptly ripped away as it reached the end of its travel.

The Board should understand that the FDR record of airspeed and altitude was not accurate from 24:30 to 24:40 FDR time, during the recovery pull-up. The dynamic effect of rapid pitch changes during this interval caused the Pitot-static system indication errors. The slat extended and immediately failed under the following conditions of overload:

_ altitude approximately 11300 feet

_ airspeed approximately 445 KCAS, Mach 0.814

_ in a near vertical dive with flight path angle of -75 degrees

_ with the aircraft experiencing a substantial right roll rate

_ approximately 4.7 G’s of normal acceleration

One of the manufacturer’s engineers wrote an account of the probable sequence of failures of the structural components of the slat. That analysis was in the manufacturer’s report, on pages B-15 and B-16.

Extend Stops sheared off the inboard slat track as the slat reached the limit of its extend travel.

A previously existing failure of the T-bolt for the inboard slat-to-track connection resulted in reduced torsional, angle of attack, rigidity of the slat. This permitted the inboard end of the slat to rotate as it reached the limit of extension.

Brackets, mounted in parallel, on the lower surface of the inner slat skin were bolted to each side of each slat track. The metallurgist examined the brackets attached to the inboard slat track:

Angle skin support outboard bracket fracture. . . reveals multiple origins along almost the entire inboard edge. . . . the slow growth region had progressed approximately 40% through the bracket wall. . . . which demonstrates a fatigue type fracture. . . . The inboard bracket . . . separations were a result of static overload. One bolt had been sheared out on the inboard angle and all upper bolt holes on both angles were severely distorted . . .[Manufacturer’s report page B-31. Emphasis added.]

The distorted bolt holes in these angle brackets for the I/B slat track are shown on page B-22 on the manufacturer’s report. The axis of the elongation in each of these holes suggests that the severe distortion may have occurred at the same moment that the Extend Stops sheared away from the I/B track. This damage to the bolt holes in the angle brackets suggests that the #7 slat reached the limit of its extend range with considerable energy. The orientation of the distortion axes suggests that as the inboard end of the slat reached the extend limit, the slat was also twisting under (inboard end twisting counterclockwise). It was this twisting that allowed the lower surface of the slat structure to contact the actuator rod. This resulted in the “static rupture” of the actuator rod near the end that connected to the slat (referred to on page B-15 of the manufacturer’s report).

The metallurgist made other observations of the #7 slat components. Those observations suggest that while the various forces pulled the slat forward from the retracted position, the air loads on the slat included a lateral span-wise component. A sideslip condition persisted as the slat was pulled from the retracted position and contributed to the transverse load on the extending slat. That transverse component of the drag loading caused the slat tracks and actuator rod to bend in an outboard direction. It also accounted for the unexpected orientation of the crease line found in the slat inner skin panel. (The reader should understand that normal extension of a slat involves both a forward and outboard component; a motion perpendicular to the wing leading edge.)

The manufacturer’s representative viewed slat parts while in the NTSB facilities in Washington D.C. His observations were in the manufacturer’s report to the NTSB. Refer to that report for each page number credited below. [Emphasis added in quotations below.]

The outboard track of the #7 slat was observed:

A view looking aft shows the track was alsobent and twisted outboard for a total outboard displacement ofabout 7.0 inches at the slat attach point.[Pg B-13.]

The inboard track of the #7 slat was observed:

A view looking aft would show track at slat attach pointdisplaced outboard about 6.0 inches .[Pg B-13.]

Theslat inner skin panel adjacent to and inboard of of the actuator rod attachment shows a crease indentation consistent with the actuator rod and housing bearing upward against the slat . . . Thecrease line extends inboard and aft from the actuator attachment point on the slatinstead of directly aft as expected for a normally positioned slat.[Pg B-15.]

Theactuator rod end fracture face alsoexhibits a slight lateral fracture direction. This is compatible with the slat sideways motion occurring prior to the rod contacting the slat.[Pg B-15.]

An air load on the slat (drag), acting aft and outboard, caused the slat tracks to deflect aft and outboard. As the tracks bent aft the slat mid section contacted the fractured actuator rod. The aft drag load on the slat caused the inner slat surface to press on that fractured end of the actuator rod. The actuator cylinder then rotated (pitched down) about its trunnion mount. The head end of the actuator then contacted the front wing spar. This created the short coupling and the bending moment that resulted in the fracture of the outer cylinder barrel of the actuator.

Then the slat box buckled and sheared apart near mid span at 3 inches outboard of the attach point for the fractured actuator rod.

The inboard half of the #7 slat then rotated aft under the air load and separated from the inboard slat track. The time interval from the beginning of the slat extension until the separation of the inboard half of the slat was probably less than one second. The forceful extension and separation of the slat was part of the “explosion” heard after the crew extended the landing gear.

The air load on outboard half of the #7 slat caused the additional bend of the outboard slat track and the crack in that track web. The T-bolt for the outboard track-to-slat connection sheared as did the other slat-to-track connect points. The outboard half of the slat separated from the aircraft about one second after the inboard half.

The records from other in-flight incidents that involved an uncommanded extension of a #2 or #7 slat documented the quickness with which it separated from the aircraft. In the documented cases the crews described their perception of the extension and separation of a slat as: “a violent shudder,” “a heavy bump,” a “sever jolt,” or “felt and heard a loud thump.” One case of slat extension and separation, EAL/N8139 on 1May69, was well documented. In that case, EAL Flight 267 was on descent to Nashville while at 325 KIAS, below 17000 feet altitude,

when the crew heard what sounded like an explosion. . . . they noticed that hydraulic pressure was lost on System “A”. . . . Upon landing at Nashville, it was noted that the entire Number 7 right wing leading edge slat was missing. It appeared during our inspection that the outboard track failed first and the slat swung under the wing at 90 degrees before breaking off. [31]

In the documented incidents involving slat separation, as in the accident of TWA 841, the slat actuator had disconnected from the slat structure. In each case the actuator rod had fractured prior to the instant the slat tracks bent aft in reaction to the drag load on the extended slat. In each case the slat tracks bent under and then the slat broke off and departed beneath the wing surface. As described in earlier paragraphs, in the accident of TWA 841 the slat actuator rod ruptured as the slat twisted. This fracture occurred just as the slat reached its limit of extension. The similarity in each case is that the slat actuator did not carry any load at that instant when the slat tracks began to bend aft. Therefore, though the specific model of actuator installed was not identical in all of the cases, the actuator rod was not a factor in the load bearing strength of the extended slat.

Proposed new sub section 2.4.7,SEQUENCE OF SEPARATION OF PARTS DURING THE RECOVERY

The following statement appeared on page A-6 of the manufacturer’s report:

trajectory analysis does not give any clear indication as to the sequence of the various parts detaching from the aircraft’s wing.

That statement was incorrect. The trajectory analysis of the fallen parts was an important tool. However, the investigators failed to utilize that tool. The trajectory analysis provided by the manufacturer served two purposes in the reconstruction of this accident. First, it served as the only semblance of a map of the trail of debris. The NTSB failed to include a proper wreckage distribution chart in the docket. Secondly, the investigator learns the separation sequence of the various parts by superimposing the parts’ trajectory over the profile track of the aircraft.

Included in the manufacturer’s report on page A-10 was Figure 4. It showed the orientation of the trail of debris. [See the graphic proposed in section 2.4.1, page 41 herein. It shows a transcription of the parts’ trajectory.] Found on the ground, spread over five thousand feet, were four parts which had ripped from the aircraft. The analyst geometrically projected the final impact points of debris onto a line oriented with the known wind direction, northeast to southwest, on a 240 degree track. The accompanying text stated that the winds below 9000 feet were from the northeast. The plot showed an erroneous label for the flap track canoe fairing found on the ground. (The Board repeated that mistake, later in the final NTSB report: the “#6 [sic] flap track [32]fairing” impacted very near the inboard half of the #7 slat.) This Figure 4 showed that the impact point of the outboard half of the #7 slat, projected to the wind direction base line, was about one thousand feet southwest of the previously mentioned parts. Plotted about thirty-seven hundred feet further along the base line was the impact point of #10 spoiler panel.

On page A-6 of the manufacturer’s report, the engineer commented about his plot of the trajectory of the separated parts: “More definitive conclusions could be drawn if the aircraft’s flight path were superimposed on the part’s trajectory.” Utilizing the flight recorder data, the analyst can fit a series of arcs together to form a curve that approximates the profile track of the aircraft during the 5 G recovery pull out. The work-up of the profile view of the track of the aircraft is shown on the next page. The pilot used the moon as the reference point for the pull up from the dive. The azimuth of the moon was 242 degrees from north. The profile track of the aircraft during the pull up from the dive roughly followed the vertical plane shown in the trajectory analysis. Therefore, employing the suggestion of the manufacturer’s engineer, the profile track of the aircraft during the pull up from the dive can be conveniently “superimposed on the part’s trajectory.” See the overlay of the plots shown on page after next. Shown is one possible result of the overlay technique. An analyst could shift the profile view of the aircraft track to the right with differing results. Such an overlay of plots yields some interesting results.

One startling result from this graphic technique is that the separation of the #6 flight spoilerapparently occurred much later in the pull up than the separation of other parts. The #10 spoiler panel is an alternative designation for the #6 flight spoiler, the most inboard flight spoiler panel on the right wing. The “B” hydraulic system, which remained pressurized, operated this most inboard flight spoiler. The Captain had positioned the speedbrake lever full aft earlier in the dive, in an attempt to deploy speedbrakes. It appears from this plot overlay that the #6 flight spoiler separated from the aircraft at slower speeds late in the recovery when spoiler blow down ceased. During the recovery, as the airspeed decreased the hydraulic forces provided by the remaining “B” hydraulic system would deploy the inboard flight spoiler panels. Acting in the speedbrake mode, the inboard flight spoiler panels would deploy from the blown down position to the extended position.

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Profile View of Recovery Pull Up

recovery profile view (FDR)

, Flight Data Recorder (FDR) parameters ,

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Overlay of Aircraft Track and Parts Trajectory

,

Separation of parts ,

The overlay of plots displays the probable point of separation of the spoiler panel. Separation occurred between Point 2 (FDR time 24:48.4, 275 KIAS) and Point 3 (FDR time 24:49.5, 260 KIAS) shown on the plot of aircraft track. This suggests that the spoiler panel separated as the airspeed decreased below 270 KIAS. The Maintenance Manual [33]states: “At 310 KIAS, blowdown is very effective . . . at 270 KIAS or less, blowdown is no longer effective. This permits all spoiler panels to extend.” This would suggest that the spoiler panel separated from the aircraft when it was forced up by the spoiler/speedbrake actuator as speed decreased below the blow down regime.

The #6 flight spoiler panel may have been initially torn during the over-speed, as was the #4 flight spoiler4 flight spoiler. Then after such weakening, it may have separated after the gear extended, due to air loads inside the opened (damaged) wing section. Increased internal pressure resulted from the damage to the large access panel outboard of the right landing gear.

One other event could have caused the separation of that spoiler panel. The separation of the #6 flight spoiler may have occurred as the pilot retracted the speedbrakes, when the airspeed was under control. At the Point 2 shown on the plot of aircraft track, the flight path angle was +33 degrees and the G forces were being relaxed; the pilot would then have considered retracting the speed brakes, if he had not already done so. If this were the case, when the pilot moved the speedbrake lever toward the retracted position the spoiler actuator would have attempted to pull the extended panel toward the retracted position. The spoiler panel may have failed during the attempted retraction, due to air loads inside the damaged wing section.

Proposed new analysis Section 2.5,THE INITIAL UPSET AND LOSS OF AIRCRAFT CONTROL

The title of the OLD Section 2.5 was:Loss of Aircraft Control . That old Section 2.5 had included twenty-one paragraphs covering various topics such as the CVR, plus a seven paragraph summary. The board had spread the old Section 2.5 over pages 27 through 33 of the NTSB Report (AAR). In the old Section 2.5 the Board included many items. Some items, such as the comments about the CVR, were unrelated to the subject of aircraft control. The Board should delete most of its stability and control analysis included in the old Section 2.5. The Board erred by utilizing incorrect assumptions as a basis of its analysis. The Board incorrectly assumed that a slat extended at FL390. The Board also incorrectly discounted sudden yaw as a factor in the upset and in its analysis of the FDR heading excursions. The erroneous assumptions spawned invalid analysis and conclusions.

Proposed newsubsection2.5.1,VIBRATION,YAW,ANDROLL

The first paragraph under the section 1.17.3 (bottom of AAR pg 18) specified a time for the onset of buzz: 2147:34 EST (23:23 FDR time). That buzz or slight vibration was noticed by the pilot. Then the pilot noticed the autopilot commanding lateral control deflections. This was apparent from motion of the control wheel. The pilot stated:

I saw the yoke, the wheel of the airplane just starting to turn slowly to the left. . . . The airplane was turning to the right. The autopilot was trying to correct. [34]

There were two possible vibration components of the initial buzz sensed by the Captain. The most probable source of vibration was the flutter of the outboard right aileron. This vibration component persisted until the landing at DTW. Secondly, as flight spoilers on the left wing began to deflect, they generated a spoiler buzz. The lateral control commands from the autopilot roll computer caused the aileron-spoiler mixer to actuate these spoilers as the aileron control wheel passed ten degrees of CCW rotation.

The #6 flight spoiler, the most inboard spoiler on the right wing, was the only spoiler panel that separated from the aircraft. There is a possibility that the #6 flight spoiler (#10 spoiler panel) actuator was defective and permitted that spoiler panel to float. Such a defect, resulting in spoiler float, is not unusual. Crews often identify such a floating spoiler panel after noting the excessive lateral control input (displacement of the aileron control wheel) required by the autopilot. Such a floating spoiler panel could have initiated the slow rolling motion to the right, thus prompting the response of the autopilot. A floating #6 flight spoiler panel could also have contributed to the lack of left roll control available to the pilot.

The onset of such spoiler float may have created the initial rolling motion which triggered a chain of interactions. The air loads from such a rolling motion may have contributed to the final overload fracture of the hinge bolt for the right outboard aileron. Over the years that bolt had developed multiple fatigue fractures. The loads on the bolt, and the fatigue progression are discussed in a new subsection 2.5.2, below. Cold soaking further weakened the bolt during that flight. The typical loads acting on the weakened hinge bolt may have been sufficient to cause the final overload fracture that propagated from the previously existing fatigue fracture faces. After that hinge bolt failed, the resulting up float of the right outboard aileron induced sudden yaw and an added rolling moment. The yaw-rate gyro in each yaw damper coupler would detect these yaw-rate changes. The manufacturer programmed each yaw damper computer to respond to yaw-rate changes. Each yaw damper computer would command sharp rudder deflections to stabilize the sudden changes in yaw-rate.

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Lateral and Directional Flight Controls

flight controls illustration , autopilot , yaw dampers , rudders , spoilers illustration ,

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Initial Right Rolling Moment

,

outboard aileron hinge bolt failure , vibration ,

Engineers included these dual yaw dampers as an engineered safety feature. The B727-100 aircraft has a natural high altitude dutch roll characteristic. This dynamic instability proved to be so severe that it could become undamped and divergent. This characteristic was an undesirable consequence of the aircraft design -- specifically the combination of wing sweep, wing dihedral, and tail volume. The engineered safety feature, the yaw damper, responded only to yaw rate inputs. Thus a yaw damper might deflect a rudder even when no dutch roll oscillations are present. Any sudden yaw-rate changes sensed by the yaw damper system could be processed into a discrepant rudder deflection.

Rudder deflections, commanded by the yaw dampers, likely sustained the nose right sideslip by opposing the restoring moment.

As the aircraft rolled to the right, the Captain “took the aircraft off the autopilot.” The Captain stated;

It, the buzz, led right into a very gentle buffeting, . . . the aircraft continued to roll to the right. And at that time I started coming in with the left rudder . . . because I was . . . to the stops on the left aileron. [35]

In the above statement, the Captain referred to an increase in the vibration intensity from a buzz to a buffet. This increase in vibration intensity may have resulted from spoiler buzz added to increased vibration intensity from the flutter of the right outboard aileron. This flutter component of vibration will be discussed later in a proposed new (sub)section 2.5.2, “Freeplay and Flutter of Right Outboard Aileron.”

The right rolling moment created by the up-float of the right outboard aileron added to the right rolling moment due to the nose right sideslip. This combination of right rolling moment exceeded the counter rolling moment generated by the pilot’s application of full left wheel throw and full left rudder pedal travel.

An anomaly known as control wheel ratcheting may have contributed to the diminished left roll authority available to the pilot. This is discussed later in a new sub section (2.5.3),titled “Autopilot Roll Channel.”

After forcing the control wheel to full left (CCW) wheel throw, the pilot moved the left rudder pedal to full travel. However the aircraft continued rolling right. The pilot’s pedal input to the rudder actuator, when summed with the signal from the yaw damper, may have produced no net left deflection of the rudder. Information from the manufacturer states that the yaw damper authority is limited to 5 degrees of rudder deflection. Either or both rudders could be driven to that 5 degree deflection limit by commands from the yaw damper coupler, signal distortion in defective connectors, or by discrepant operation of the electro-hydraulic servo valve (in the rudder PCU).

After the aircraft began the yawing motion to the right, the yaw rate gyro would detect any change in that yaw rate. It is possible that, as the pilot pushed on the left rudder pedal, each yaw damper sensed the change in yaw rate, and commanded its rudder to deflect right. The yaw dampers only sense yaw rate, not sideslip. Once established in a nose right sideslip, the natural restoring moment caused by the vertical fin would produce a nose left yawing moment. The yaw dampers’ signals to the rudders might work to counter this restoring moment, by acting to oppose a reversal of the existing nose right yaw rate. That is the logic by which yaw dampers resist the dutch roll oscillations, and that seems to have contributed to some other high altitude upsets. Thus, after the fault in the lateral control system, the interaction of the automatic directional control system could actually work to sustain the existing nose right sideslip.

Some factors that influenced the lateral and directional stability of the accident aircraft are illustrated on the next five pages.

_

Intentionally left blank.

Yaw Damper Interaction

,

yaw damper, rudder deflection , FDR Heading changes

Intentionally left blank.

Negative Sideslip Angle Caused By Yaw

sideslip illustration , parameter relationships , yawing moment

Intentionally left blank.

Vertical Stabilizer and Rudders

,

rudder-fin illustration , sideslip , vertical stabilizer ,

Intentionally left blank.

Sweepback and Dihedral Effects

,

rolling moment , sweepback effect , dihedral effect , sideslip

Intentionally left blank.

Left Wing Flight Spoilers Extend -- Spoiler Buzz

vibration , spoiler buzz , autopilot lateral inputs ,

Continuing with his description of the upset, the Captain testified,

I had full aileron input and full left rudder input also, and the aircraft was rolling to the right. . . . going through about thirty degrees. . . . And it continued over . . . I got the throttles closed and grabbed back onto the yoke again because . . . you could feel the pressure on the yoke . . . from the control inputs. [36]

The cause of such a yaw induced rolling motion is difficult for the pilot to identify quickly. At night pilots are not likely to perceive yawing motion, especially when distracted by other tasks (such as stowing charts). Perception of a control problem would be delayed until the pilot sensed the rolling motion. In this case, the upset was sustained by a complex interaction involving lateral and directional controls. Unknown to the flight crew, the high altitude and gross weight were a critical combination for that model of aircraft. An otherwise minor control problem could become threatening under those particular conditions, due to the tightly coupled interaction of yaw and roll.

Once disturbed the left wing rose, and for reasons unknown the normal restoring moment was inhibited. With full left wheel throw, there was not enough lateral authority from ailerons and flight spoilers to level the wings. As the right wing continued to drop, a yawing moment persisted that caused the aircraft to turn to the right. Full left rudder pedal input by the pilot did not yield sufficient rudder deflection to stop the sideslip induced rolling motion. As the right bank increased, the vertical component of lift was not enough to counter the weight and the aircraft lost altitude. The aircraft started into an ever tightening spiral dive.[37]The aircraft suddenly became controllable again only after the “A” System Hydraulics failed as the right gear extended.

Intentionally left blank.

Yaw Damper Fail Flag

,

annunciation ,

The cockpit lacked any annunciation for yaw damper faults except for a failure of electrical power to the yaw damper coupler. In the TWA configuration, the yaw damper fail flag monitors only electrical power. The lower yaw damper fail flag, noticed by the crew following the recovery, indicated only that there had been electrical power interruption to the lower yaw damper system. The NTSB stated that after the flight the “yaw dampers” were removed from the aircraft and functionally tested. That bench check, of the yaw damper couplers, identified no discrepancies. The NTSB erred when it failed to explain the cause or the possible effects of that yaw damper failure annunciation. And the NTSB erred by failing to impound both rudder actuators, and both yaw damper transfer valves,that had been removed and replaced following the accident. As the CAB learned during its investigation of the AA Flight One accident, a thorough tear-down analysis of the rudder actuators can reveal defects not identified by the manufacturer’s “functional tests.”

In an effort to analyze this high altitude upset, the analyst should consider one other factor that may have caused the yaw dampers to command full right rudder: possible gimbal errors in the rate gyros of each yaw damper coupler. The FDR showed an abrupt change in the heading trace at FDR time 23:42.8, from a right turn to a left turn. The NTSB attributed this reversal to gyro gimbal errors in the #1 directional gyro (DG). Had the yaw rate gyros experienced an error similar to that of the #1 DG, each yaw rate gyro may have suddenly signaled a left yaw. This input may have triggered each yaw damper to command right rudder. It is possible that those rate gyros experienced similar gimbal error even earlier than did the #1 DG, before the FDR time 23:42.8. Thus, such yaw damper inputs could have occurred while the aircraft was still at 39000 feet, further complicating the initial upset.

The rudder control systems are critical to the safety of B727 aircraft operating at high altitudes. However, as installed on the accident aircraft, those systems are poorly instrumented and lack any means of fault annunciation. Small indicators positioned at the bottom of the instrument panel are easily overlooked or misinterpreted. Indicators are more difficult to interpret when small pointers are active, and travel over a short range. The scale of the instrument may be fine for some flight regimes, but not practical (or unusable) during another phase of flight.

Intentionally left blank.

Rudder Position Indicator

,

rudder hard-over indication , Rudder Position Indicator illustration

Such is the case of the rudder position indicators and the yaw damper fail flags: size, position, and scale make the information displayed less accessible. Illustrations show the Rudder Position Indicator that was installed on the B727 involved in the ’79 yaw roll-over accident. As the illustrations on the previous page depict, the indicator scale is not usable during the cruise condition. Full rudder pedal travel yields only a small movement of the pointer, only to the first marking on the instrument. In contrast, when a pilot performs the flight controls check prior to takeoff, full pedal travel yields full scale deflection of the pointers.

The practical accuracy of the rudder position pointers can be shown to be affected by static electricity. Simply cleaning (rubbing) the glass face of the instrument can cause the pointers to deflect randomly to either side of neutral moving about the same distance across the scale as occurred during the flight test (full rudder pedal input under cruise conditions). So during the cruise conditions, due to the poor accuracy of the pointers and the cramped scaling of the instrument, the Rudder Position Indicators are a poor tool for evaluating any sort of rudder or yaw damper malfunction.

Proposed new sub section2.5.2,FREEPLAY AND FLUTTER OF THE RIGHT OUTBOARD AILERON

As part of its analysis, the Board discussed this freeplay under the old section 2.4 in the 16th paragraph (second paragraph from the bottom on AAR page 26). In its analysis, the Board did not consider the flutter of that freed outboard aileron. The Board did not consider complex interactions that resulted from the sudden yaw and roll induced by uncommanded deflection of the outboard aileron. For that short body B727, at 39000 feet, there existed a tight coupling of the lateral and directional stability. Under such conditions, at night, the pilots were dependent upon the accurate response of the engineered safety features, the yaw dampers.

The Structures Group Report, dated May 10, 1979, stated:

There was 1 inch up and 3/32 inch down free play at the outboard aileron trailing edge as measured from the faired position. This was measured with the aileron locked out. It was found that the mid bolt at the actuator rib had fatigued and the other holes were sloppy.(See metallurgy laboratory report.) The nut end of the bolt had backed out outboard and remained captive. The head end had backed out of the hole and was missing.[Emphasis added.]

The NTSB metallurgist’s report (dated August 23, 1979; page 5, fourth paragraph) stated that the failed bolt

had markings indicative of fatigue progression over most of the fractured cross section. Detailed SEM examination disclosed that two fatigue cracks had propagated toward each other.. . . One of the cracks was relatively large, initiating from multiple origins along the outside diameter surface with fatigue progressive in a near transverse plane. The other crack was much smaller, also initiating from multiple sites but with propagation mainly on an inclined plane. Only a small amount of the fracture in the area where the two cracks came together was typical of an overload break. . . . The immediate origin areas for both of the cracks were covered with deposits believed to be corrosion products.[Emphasis added.]

Under typical cruise conditions, the resultant transverse load on that aileron hinge mid rib bolt was 1124 Lbs. That load was a fraction of the ultimate bolt strength of 14600 Lbs. [That load calculation is based on data shown for bolt R Bin Figure 4,727 Outboard Aileron Capability and Freeplay with Broken Bolt . . .; under cover letter dated April 16, 1980; Boeing’s Hogue to the NTSB’s Kampschror.]

Possible freeplay flutter was discounted in a letter (part of the docket) from Boeing/Hogue to NTSB/Kampschror dated Dec 3, ’79:

The outboard aileron is mass-balanced and consequently has an adequate flutter margin. This margin has been demonstrated by wind tunnel testing to be in excess of 1.2 times the dive speed (V D) of the airplane.

At a presentation on March 18, 1980, the manufacturer provided the NTSB’s investigators with analysis that discounted the effects of a free floating outboard aileron. In the written comments included in the NTSB docket, the manufacturer focused on possible effects on the buffet characteristics of the aircraft. In that particular analysis, the manufacturer did not consider possible flutter of a free floating outboard aileron. However, manufacturer’s data suggested that the free-floating outboard aileron would be driven into a dynamic response. The manufacturer described the data:

Figure 4 . . . shows the pressure distribution on a 727 wing section at a typical cruise angle of attack and Mach. Also shown is the pressure distribution resulting when the aileron is deflected upward one inch at the trailing edge. The data of the figure show . . . that as the aileron is deflected upward the load on the aileron changes direction. The pressures on the lower surface become less than the pressures on the upper surface. This causes an aerodynamic down load to develop on the aileron that tends to drive the aileron back toward the faired position. [38][Emphasis added.]

That weakened hinge bolt for the right outboard aileron probably failed at 2147:34 EST (23:23 FDR time).

VIBRATION

The last paragraph on page 23 of the NTSB’s AAR stated:

the Safety Board concludes that an extended No. 7 leading edge slat . . . created a buzzing noise or slight buffet followed by moderate buffet.

The Board’s vibration analysis of the “buzz” or “buffet” was incomplete. The comments from the witnesses indicate that the vibration had several components. Vibration marked the very beginning of the event and continued until the aircraft landed at DTW. The vibration persisted even after the #7 slat had separated from the aircraft.

At the beginning of the incident, crew members noted the initial buzz that changed to a light buffet. [39]Passengers also commented on the initial vibration:

Nine passengers related the initial vibration or shaking to various degrees of turbulence. However, one passenger described it as strange because everything was too perfect and the ripple effect was not random. (“It felt like a perfect distance between pulses and the same amplitude every time.”) This very sharp, choppy feeling continued 5 to 10 seconds. [40]

After the recovery, when TWA 841 contacted Cleveland Center, one pilot transmitted,

level at ten thousand now under control we got a severe vibration . . . [ATC transcript, 0250:00 GMT.]

Passengers also commented on the persistent vibration:

After apparent regain of control, the vibration changed but continued until landing. The vibration periodicity remained about the same but the intensity was substantially diminished. [41]

The “A” Flight Attendant reported:

The aircraft started to vibrate unusually. . . . We pulled out of the fall . . . . Vibration of the aircraft continued. [42]

From the Captain’s testimony:

Q After your pull-out, did you notice any damage to the aircraft?

A Yeah, it shook, vibrated; it buffeted severely, made an awful lot of noise, awful lot of buffeting noise. [43]

The crew accomplished a series of emergency checklists due to the loss of System “A” Hydraulics. The pilots had extended the landing gear by normal means during the uncontrolled dive. However, multiple gear unsafe indications required the crew to accomplish the Manual Gear Extension Checklist. The Captain testified:

The buffet subsided to a certain degree when we went through nose gear extension. [44]

The “B” Flight Attendant reported:

I noticed a definite vibration that was to continue until the end of the flight. I knew it wasn’t turbulence. It was something I’d never experienced before. . . . After . . . a period . . . we began to level off but there was still the vibration, especially strong up by the forward F/C lav. [45]

Due to the partially extended nose gear and associated damage, the position of the mechanically linked nose gear doors caused aerodynamic buffeting in the nose wheel well. That vibration component diminished after the crew executed the manual extension procedure for the nose gear, which repositioned the doors. It was flutter of the right outboard aileron that was the one continuous vibration component. That source of vibration lasted from the beginning of the uncontrollable flight segment, until the landing at DTW.

At various instants after its initial onset, this vibration sensed by the witnesses was the sum of several components:

___flutter of the right outboard aileron

___buzz from the flight spoilers responding to autopilot commands

___high speed buffet

___buffet in the nose wheel well until after manual gear extension

The NTSB’s examination of the G-trace, from the accident FDR, showed that slight vibration began at 23:23 FDR time (2147:34 EST).

The NTSB’s examination of the Heading trace from the accident FDR showed the onset of a right yawing motion at 23:30.2 FDR time (2147:41.2 EST). The Autopilot Roll Computer and the FDR Heading stylus received this heading change signal from the #1 DG. There was a seven second lag between the onset of vibration and the onset of the heading change. The heading change would have triggered the Autopilot to respond with lateral control inputs only after 23:30.2 FDR time. Thus the deflecting flight spoilers on the left wing could have added a vibration component (due to buzz) only after 23:30.2 FDR time.

Tension of the FDR Heading stylus against the foil may have contributed to the time delay of the recorded heading change. Post accident tests showed that FDR stylus tension was one variable that affected accuracy of the measurements recorded on the FDR foil. [46]That delay in the movement of the stylus, of undetermined duration, likely resulted in the instantaneous heading change recorded on the FDR foil. Actual aircraft heading changes, properly sensed by the autopilot, may have preceded the initial heading change recorded by the FDR. However, it is unlikely that the autopilot sensed this heading change more than a second or two prior to the movement of the FDR heading stylus.

That initial buzz was a vibration induced by the onset of flutter of the right outboard aileron. This is a deduction that grows from the fact that there was the seven second time difference between the onset of vibration and the first recorded heading change. The pilots had established the accident aircraft in the heading hold function of the autopilot’s manual mode. During the initial upset any autopilot roll commands were a response to a perceived deviation from the reference heading. Airplane heading deviation would then cause the autopilot roll channel to command proportional bank angle in the proper direction to regain the reference heading. [47]That seven second interval, between vibration and heading change, precluded the possibility that normally deflected flight spoilers, responding to autopilot commands, were the source of the initial buzz.

Proposed new sub section2.5.3,AUTOPILOT ROLL CHANNEL

Under the old section titledLoss of Aircraft Control , on pages 27 - 33, the Board did not consider the effects of the control wheel ratcheting. The Board should discuss the results ofthe 1977 testing done by TWA , aboard the very same aircraft -- N840TW -- that was later involved in this accident.

Earlier in this petition we presented factual information concerning that 1977 testing, in a proposed new (sub) Section 1.16.5, “TWA Tests Aboard N840TW.”

As was described in the affidavit regarding the 1977 testing at TWA, there was a major control problem that resulted from the failure of the autopilot disconnect function. That precipitated the control wheel ratcheting which resulted in diminished roll control available to the pilots. Sustained, uncommanded roll control inputs were also a factor.

The current TWA B727 Flight Handbook (FHB) Systems Description portion of the Autopilot section has a paragraph that maypartiallydescribe this anomaly. The following information is printed on FHB Page 9.01.04, column two, first paragraph.

The autopilot will not disengage if it is overpowered, but the following action may result: Overpowering the aileron may cause ratcheting of a clutch which will result in the aileron control wheel getting out of synchronization with the autopilot. When this occurs the autopilot must be disconnected to return the aileron wheel to a normal position. Overriding the autopilot is not recommended.[Emphasis added.]

This excerpt from the TWA FHB bears unique wording, and suggests an aspect of an autopilot false disconnect not found in manuals of other airlines. The anomaly that is hinted at in the above excerpted paragraph pertains to any attempt to disconnect the autopilot during an abnormal condition. The effect of this anomaly could prove critical if the pilot were to rotate the control wheel, in an instinctive struggle to stabilize an upset aircraft, while using his thumb to depress the Autopilot Disconnect Switch. Two new problems might result that would further complicate the instability:

1) The autopilot may not actually disconnect even though the paddles might trip to theDISENGAGEDposition. If the autopilot clutches remained engaged and continued to exert forces on the lateral controls, the pilot would experience heavy stick forces and uncommanded control inputs.

2) The control wheel might “ratchet” to an un-centered neutral such that the pilot’s attempts to center the control wheel would result in additional unexpected rolling moments. The new un-centered neutral point for the control wheel might further confuse the pilots by providing erroneous feedback about the position of the ailerons. Such an un-centered neutral point may result in reduced lateral control. Full wheel throw in the direction of previous ratcheting may yield only a fraction of the normal rolling moment due to aileron. Under the abnormal initial conditions that could provoke this wheel ratcheting effect, it would be very difficult for the pilot to quickly determine the control inputs needed to stabilize the aircraft.

The fact that the TWA FHB mentions the false disconnect aspect of this problem, and that this aspect was not found in manuals of other operators, suggests that TWA may have been aware of this problem as a result of the tests conducted on aircraft N840TW in 1977. It is possible that TWA had documented the false disconnect anomaly even before the 1979 – 1981 investigation of the TWA841 accident, yet failed to inform the NTSB investigators about the anomaly. There was nothing in the NTSB docket that mentioned the results of the 1977 test incident at TWA aboard N840TW.

The Board should consider the autopilot roll channel as a contributing factor in the initial upset at FL390. The autopilot roll channel may also have induced the later upset that occurred when the crew attempted to accomplish the Alternate Flap Extension Checklist (required as part of the hydraulic failure procedures). Two such lateral upsets with the autopilot engaged during the same flight suggest that a common anomaly may have induced both upsets. (The lower rudder, “A” System actuator, would have been de-powered during the second upset. Therefore that--and the lower yaw damper-- can be excluded as contributing to the second upset.) The autopilot roll channel excursions could have been induced by a faulty #1 Directional Gyro (DG). A faulty #1 DG may have sent erroneous heading signals to the autopilot roll channel and the FDR Heading stylus. The autopilot roll channel would attempt to chase a drifting #1 DG and command unexpected roll inputs. Such an intermittently drifting #1 DG was suspected as a factor that contributed to a flight control incident on December 16, 1989, near Dayton, Ohio.

Proposed new analysis Section 2.6,FAULT ANALYSIS OF SLAT RETRACT LOCKING MECHANISM

In contradiction to the NTSB’s assertion in line 8 of the last paragraph on page 24 of the AAR, direct evidence suggests that the #7 slat did not extend until after the failure of the “A” Hydraulic System. A major mistake in the NTSB’s analysis was their assumption that an extended #7 slat had caused the initial upset. The Board must recognize that the extension of the #7 slat occurred later and was a consequence of loss of hydraulic pressure to a faulty slat actuator.

The Board can exclude the possibility of a transverse fracture of the actuator piston because the slat extended only after decay of retract side hydraulic pressure.

Section D of the manufacturer’s report dealt with the failure analysis of the leading edge slat actuator. However, that failure analysis was incomplete. The manufacturer provided the Board with only limited information. The scant fault analysis tended to substantiate the manufacturer’s assertion that no reasonable failure of the retract locking mechanism could result in the actuator unlocking from the retracted position.

The information presented separately in Boeing Operations Manual Bulletin 75-7 [48]described specific component damage within the slat actuator that had developed during usage. Such faults could result in the slat being pulled from the retracted position by air loads after a hydraulic failure. In contrast, the failure analysis presented in the manufacturer’s report was only a theoretical exercise and did not reflect actual test results nor in-service failures.

The Board should acknowledge the following deduction: Since the #7 slat extended only after the failure of the “A” System Hydraulics, the slat actuator mechanical retract lock either suddenly failed or had never locked.

If the slat actuator had not locked in the retracted position then the Retract Lock Indicating Switch should have tripped the cockpit annunciation. Thus the amber, #7 slat In Transit annunciator (on the F/E’s Auxiliary Panel) and the amber “LE Flap” annunciator (on the pilots’ Center Instrument Panel) should both have illuminated. [49]A faulty Retract Lock Indicating Switch, failed in the “locked” position, could not actuate the amber light for the affected slat. The cockpit in-transit lights would then remain extinguished regardless of slat position (manufacturer’s report page D-10).

Such a failure in the Retract Lock Indicating Switch could easily go unnoticed as long as other components were normal. The indications for that slat would be normal when it extended (green light illuminated), and seem normal when not extended (no amber light, no green light). Such a fault would disable the amber in-transit annunciator only for that specific slat. With otherwise normal hydraulics, a flight engineer could only identify this failure after unusually close observance. During the brief intervals as the slats retracted, the F/E would need to face aft to gaze at the F/E’s Auxiliary Panel. He would then need to note the absence of the amber light for that specific slat. However, during line operations such a lack of in-transit indication has often gone undetected. During the F/E’s training, instructors emphasize many other duties. Each of those small in-transit indicators would illuminate for only a brief interval. A F/E may often miss the chance to verify the amber in-transit light for each of the fourteen leading edge devices. There was no procedural requirement that obligated a F/E to verify the proper in-transit annunciation of every leading edge device.

Most likely, accident investigators had difficulty understanding the Slat Actuator Retract Lock Mechanism. Because of numerous errors in the records relating to the Retract Lock Indicating Switch, text and illustrations found in the NTSB’s docket were misleading. Even the accurate illustrations were of little use. Most illustration had a compressed scale. Therefore, the illustrations failed to clarify the complex orientation of the locking springs, locking pistons, and keys -- all inside the slat actuator piston.

However, an analyst can utilize several means to gain understanding of the retract locking mechanism in the slat actuator. The description given on pages D-2 and D-3, and the illustration (Figure 3) shown on page D-18, of the manufacturers report, can be of some use. By ordering the color photographs from the NTSB docket, the reader can receive two photographs (at different levels of magnification) of a disassembled slat actuator cylinder, piston, and rod. The best aid, in understanding the retract locking mechanism of the slat actuator, is to view and handle the components. By viewing and handling a cutaway slice of an actuator piston, a researcher can understand its internal orifices. Through these orifices the hydraulic forces act, the Locking Pistons and Springs travel, and the Lock Keys slide.

POSSIBLE FAILURE OF RETRACT LOCK KEYS TO ENGAGE

Page D-11 of the manufacturer’s report described the effect of a failure involving the slat actuator Retract Lock Keys. In this “event” the Retract Lock Keys do not extend into the locked position behind the Retract Lock Ring. The manufacturer identified only one cause for this “event”: excessive friction in the lock mechanism. Other causes could also result in exactly the same “event” in which the Lock Keys fail to extend to the locked position. For example, the misaligned #7 slat might hydraulically retract fully against the outboard retract Hook Stop. Due to the slat misalignment, the actuator rod might inhibit the actuator piston from retracting far enough into the head end of the cylinder. Thus, the retract Lock Keys may not engage behind the Retract Lock Ring. In this case, only the hydraulic pressure against the retract side of the actuator piston would ensure that the slat remained in the retracted position. When hydraulic pressure failed there would be no back-up mechanical lock to retain the slat in the retracted position.

The Metallurgists Factual Report, dated 23Aug79, documented misalignment of the #7 slat. On page 4 he reported that the outboard Hook, mounted on the #7 slat, showed

notable wear as if the hook had been engaged in the hook stop. There was also a deformation to the hook tip suggestive of impact damage. . . . The inboard hook . . . still contained its zinc chromate surface finish and showed no signs of wear with the mating hook stop in the wing.

The Structures Group Report stated that the Extend Stops for the inboard track had sheared their bolts. This record of wear and damage indicated that the #7 slat had been poorly aligned. While in the retracted position the inboard end of the slat did not properly nest. This slight gapping may have been of sufficient magnitude to have inhibited the travel of the actuator piston and thus prevented the Locking Keys from engaging.

The manufacturer stated the effect of such a failure of the retract Lock Keys to engage:

The actuator would remain in the retracted position until commanded to extend due to system pressure acting on the retract side of the actuator. Ade-pressurized actuator would be free to move dependent upon external slat loads. The“Leading Edge Flap” amber light to the affected slat would remain on when slat is retracted or extending.[Emphasis added. Manufacturer’s Report, Page D-11.]

The next analysis section discusses the possible malfunction of the Retract Lock Indicating Switch. Such a fault could disable the “L E Flap” amber light for the #7 slat so that the crew would receive no slat unlocked annunciation.

POSSIBLE FAILURE OR ABSENCE OF THE RETRACT LOCK PISTON SPRING

If the #7 slat had properly locked in the retracted position a fault could have permitted the retract Lock Keys to disengage following the loss of hydraulic pressure. Page D-14 of the manufacturer’s report stated:

Event: The lock piston spring fails or is inadvertently left out during actuator assembly and ‘A’ system [hydraulic] pressure is lost.

Result: No spring load or a reduced spring load would be available to hold the lock piston in the locked position and pressure would be lost on both sides of the piston.

Effect: The actuator would remain locked. Lock piston seal friction and friction between the lock piston and lock keys would be sufficient to retain the lock piston, regardless of external load applied to the rod.

The slat actuator piston departed the accident aircraft and was never found. Investigators were therefore uncertain about the condition of the Retract Lock Piston Spring for the #7 slat actuator.

The “effect” of a failed or omitted Lock Piston Spring as quoted in the fault analysis may be erroneous. The failure analysis did not consider the specific conditions that existed during the uncontrollable flight segment of the accident, as the right landing gear over extended. A variety of forces may have excited the locking mechanism of the slat actuator. The sudden hydraulic pressure surges associated with the gear extension may have disturbed the Retract Lock Piston and Retract Lock Keys. Following the loss of hydraulic pressure the effects of the severe vibration probably would overcome friction resistance and dislodge the Retract Lock Piston and the Retract Lock Keys. High G forces may also have contributed to dislocation of the Lock Piston and Keys from a previously locked position. Additionally, force from the Retract Lock Switch acting through the Switch Lever and the Sensing Pin may have affected one Retract Lock Key.

These various forces – especially vibration – acting on the lock keys after failure of the hydraulic pressure would likely exceed any friction forces. The worn seals of an old slat actuator could not provide the same tightness and friction forces of newer seals. The failure analysis unrealistically over estimated the reliability of friction forces to restrain the moving parts of the slat actuator.

Ronson Corporation conceded such a dominance of vibration forces over friction forces in the absence of any hydraulic forces. The NTSB documented this in a letter from L. D. Kampschror (NTSB) to McIntyre (ALPA), dated May 1, 1980.

Ronson agrees that vibration of the actuator probably would cause the retract lock portion of the fractured piston to unlock. Therefore, we find no need to test an actuator to prove these points. [Emphasis added.]

The subject of the Kampschror letter was a failure resulting from a fractured main piston in the slat actuator. Yet, the words regarding the vibration forces are applicable to other failure modes involving hydraulic pressure loss and a failed or missing Retract Lock Piston Spring.

POSSIBLE FAILED SEAL AROUND THE

SENSING PIN FOR THE RETRACT SWITCH

The Board should consider one other failure “event” mentioned in the manufacturer’s report on page D-6. It described a failed seal at the retract switch Sensing Pin. This fault, of a worn seal, would cause the cavity of the retract switch to eventually fill with fluid. This skydrol might then leak overboard from the switch housing. Such a fault and resulting leakage may explain the skydrol bathing found on the right wing upper surface of the accident aircraft.

The manufacturer did not clearly state the second “effect” of such a seal failure (last paragraph on page D-6 of the manufacturer’s report). The hydraulic fluid would pressurize the switch cavity during each slat extension. When hydraulic fluid ported to the head end of the actuator cylinder it would leak past the worn seal into the cavity of the Retract Lock Switch. This might cause the Retract Position Indicating Switch to remain in the “locked” position. The amber slat In Transit annunciation would thereby remain extinguished even when the slat had not locked in the retracted position. Contrary to the wording in the manufacturer’s analysis, the slat actuator would extend or retract in response to the hydraulic forces (if available).

In the TWA 841 accident, hydraulic fluid may have leaked past a worn seal at the retract switch Sensing Pin for the #7 slat. The mechanical retract lock may then have failed to engage after slat retraction due to the documented misalignment of the slat. That combination of faults could have co-existed passively. The subsequent loss of hydraulic pressure then permitted an uncommanded extension of the #7 slat. The evidence of the skydrol bathing on the upper wing surface aft of the #7 slat supports this failure sequence. This failure sequence also explains the lack of the amber in-transit light either before or after the slat separated. Only such a compound failure in the slat actuator could permit the #7 slat to immediately extend, and rip away, when the hydraulics failed.

Proposed new analysisSection 2.7 ,FAULT ANALYSIS OF SLAT RETRACT LOCK INDICATING SWITCH

Leakage from the area of the slat actuator may have emitted from the indicating switch. Worn seals in the actuator-to-switch interface may have caused the switch failure due to immersion and viscous friction, and the leakage of skydrol (bathing stain) across the upper wing surface.

A revised NTSB Report should include a diagram of the Retract Lock Indicating Switch and Slat Locking Mechanism as roughly shown in the Boeing Report on pg D-18. It is important that the components of the locking mechanism be identified.

The crew reported NO INDICATIONS of any slat abnormalities or slat warning lights after the dive recovery. In post accident testimony, the crew specifically stated that after the dive recovery the failure annunciations included: loss of the main hydraulic system, a lower yaw damper flag, and red gear warning lights. The amber leading edge annunciation did not illuminate. [50]The amber slat annunciator, located on the pilots center instrument panel, should have illuminated after the #7 slat unlocked and separated from the aircraft. The fact that that annunciator may never have illuminated is a key point in the fault analysis of the retract locking mechanism and the associated switch that activates both amber slat annunciators. There should have been a spring force within the switch housing that should have repositioned the switch after the slat unlocked. Normally this switch spring force acts against one of the retract lock keys through the lever/pivot pin, and the sensing pin. The important deduction here is that there seems to have been a fault within the retract locking mechanism that related to this switching irregularity.

Proposed new analysis Section 2.8,INVESTIGATIVE ERRORS SPARKED ERRONEOUS FINDINGS

Investigators fixated on the separated slat and assumed that the slat extended while the aircraft was at FL390. The Board even acknowledged this investigative fixation, writing in their response to a petition:

. . . the greatest portion of the Safety Board’s investigative efforts were expended in an attempt to identify any condition which might have permitted or caused the independent extension of a single slat. [51]

In the last paragraph on page 26 of the AAR, the Board focused on the erroneous premise that the slat extended while the aircraft was at FL390. Evidence appeared contradictory. The Board concluded that an unscheduled extension of the slat was improbable. Yet, the crew denied that slats could have extended as a result of their actions. The Board could have resolved the apparent conflict in this evidence by considering the possibility that the slat had remained in the retracted position until late in the dive. However, the manufacturer easily dissuaded the Board each time that NTSB investigators inquired about other possible causes of the accident. The manufacturer successfully persuaded the Board that the crew was the only possible cause of the accident.

Such a false assumption, led investigators to extensive use of circumstantial evidence. With the help of the manufacturer, NTSB investigators labored to substantiate their incorrect hypothesis.

SYNTHESIZED PITCH AND ROLL HISTORY

Based on heading gyro tests and simulation, the Board synthesized a time history of bank-angle and pitch attitude. The model included a right rolling motion followed by a left rolling motion, then a sustained right rolling motion. The Board associated each of these rolling motions with precise times. (See the seventh paragraph under section 2.5; second from the bottom on AAR page 28. The Board also presented the calculations, in graphic form, on AAR page 29.) The Board assumed that the sudden indicated FDR Heading changes were invalid. The Board assumed that gimbal error induced these indicated heading changes due to roll-rate or bank angle. The Board discounted the possibility that the aircraft actually sustained several sudden yaw motions indicated on the FDR Heading trace (AAR page 28, line 4).

Though their history of rolling motions seems probable, the Board’s basic assumption may be wrong. If applied outside the bank-angle derivation, that assumption would lead an analyst to ignore the problem of yaw induced rolling motions.

In estimating the bank-angle history during the upset, the Board recognized the unreliability of the FDR Heading trace. The bank-angle synthesis process seemed to be a trial and error (iterative) effort. The Board assumed that the sudden yaw impulses indicated on the FDR heading trace were instead gimbal errors that had been induced by roll-rate or bank-angle. Then the Board attempted to define bank angle data that would have induced those heading “errors.” The Board erred when it effectively replaced the actual FDR heading data with its own calculated bank angle data. The Board then based its instability analysis on the synthesized bank-angle calculations.

Bank angle calculations, as synthesized by the Board, may not have accurately modeled the rolling motions of the accident aircraft. The Captain testified, “. . . it went maybe ten degrees off to the right [of] this heading before it rolled.” [52]This statement, and studies of other mishaps, suggest that the Board erred when it discounted sudden yaw as the cause of the heading changes. In fact, yaw induced sideslip was a major factor in other accidents.

Perhaps an analyst could utilize short intervals of the Board’s calculated bank-angle data. However, any resulting analysis would still be suspect. (The period of the roll rate reversals suggests an underlying decay in damping. That could have resulted from a passive, if not active failure of a yaw damper.)

EVIDENCE SELECTION AND ANALYSIS

Assisted by the manufacturer, the NTSB investigators accumulated massive quantities of circumstantial evidence during the investigation of the TWA 841 accident. Investigators also noted some direct evidence from the accident. However, in cataloging the direct evidence the NTSB investigators misreported aircraft damage, failed to utilize a wreckage distribution chart, and mostly rejected the testimony of flight crew members.

During its investigation the Board focused on selected circumstantial evidence created at the manufacturer’s facilities after the accident. The NTSB erroneously assumed that an extended slat had caused the upset. That assumption tainted all phases of their investigation, and motivated the Board to attempt to substantiate their erroneous hypothesis with creative analysis of selected circumstantial evidence. The Board never attempted to correlate the direct evidence — the trail of debris, the specific damage to the accident aircraft, and the testimony of the witnesses.

In writing the NTSB Report, the Board created many pages of analysis relating the initial upset of the accident aircraft to selected test conditions performed during simulator trials and during the October 1980 flight test. The Board focused on data accumulated during selected test conditions that involved extended Leading Edge Slats and Trailing Edge Flaps. The Board’s analysis of this selected test data became the foundation of their official Aircraft Accident Report and the later NTSB Response To Petition For Reconsideration. The Board’s analysis of the selected test data also served as the core of the controversy surrounding the accident investigation, and distracted all parties away from any attempt to correlate the direct evidence. The dispute over the Board’s analysis of the selected circumstantial evidence not only taxed the resources of the NTSB, but also totally absorbed the limited technical resources of the ALPA in their effort to defend the pilots against the NTSB’s accusations. (Because these parties to the official investigation were preoccupied with analysis of the circumstantial evidence, the burden of objective correlation of the direct evidence was taken up by a private citizen; as described in the proposed section 1.19 herein.)

Much of the disputed analysis included in the NTSB’s Aircraft Accident Report and Response To Petition For Reconsideration (of Dec 1983) involved comparison of FDR traces from the accident aircraft with recorder traces created during the October 1980 flight test. The Board discussed its analysis of the airspeed trace comparison on AAR pages 13, 23 and 27; and in the Response pages 5-9. The Board discussed its analysis of the g-trace (vertical acceleration) frequency comparison on AAR pages 13, 22, 23, and 28; and in the Response pages 9-10.

Using circumstantial evidence and incomplete methods of analysis, the Board attempted to show that the pilots had extended TE flaps and LE slats while at FL390. The Board used that analysis to justify their assumption that an isolated extended slat had caused the sustained loss of control suffered by the accident aircraft. The direct evidence conflicted with the Board’s analysis.

___The pilots testified that they had never even considered the extension of flaps during high altitude cruise.

___The manufacturer’s structural engineers had admitted to the NTSB that an extended slat could not have sustained loads beyond the first third of the uncontrolled flight segment. [Manufacturer’s Report to the NTSB, page B-5.]

___The Human Factors Group surveyed passengers. During the quiet cruise condition prior to the upset, none of the passengers had sensed any disturbance associated with the loud flap motors on the B727. Yet the NTSB’s hypothesis implied that the flap motors had operated for several seconds to drive the TE Flaps to near the 2˚ position, thereby triggering the extension of the LE slats 2-3 and 6-7.

___Had any slats been extended during the cruise condition, tests show that “moderate buffet” [53]would result. Passengers and crew testified that the intensity of the vibration, associated with the initial upset of the accident aircraft, was only a “buzz” and then “a very gentle buffeting.” [54]That low intensity of vibration was not symptomatic of extended slats. [55]

Rather than re-argue the disputed points of the NTSB’s analysis of the circumstantial evidence, we ask that the Board consider the direct evidence. The direct evidence suggests that most of the Board’s analysis was irrelevant, because an extended slat could not have caused the upset and sustained loss of control experienced by TWA 841.

SIDESLIP EFFECT ON LATERAL AND DIRECTIONAL STABILITY

The Board discussed sideslip on pages 30, 31, 32 and 34, of the AAR. The Board stated:

In the B-727 full deflection of lateral controls and full deflection of rudder in the same direction can produce sideslip angles of 4.5 _to 6.5 _. Also, flight tests in conditions similar to Flight 841’s at 39,000 feet showed that rolls with full deflection of the lateral controls produced sideslip angles of about 5 __in the direction opposite to the roll. [56]

That statement revealed data not presented in the Factual Information section of the NTSB report. Neither was it presented in the records in the NTSB’s docket, in which flight test data plots showed sideslip measured only as differential pressure in inches of H 2O. The reader remained uncertain about the specific origin of that important information from the Board’s Report.

The Board’s consideration of sideslip effects focused on the case of an aircraft with an extended slat. Thus their analysis of the limited sideslip data and the associated control margin was not pertinent to the initial conditions of the accident aircraft.

In a revised report on the TWA 841 accident, the Board will face the difficult task of discussing some de-stabilizing effects of sideslip. That discussion should carefully deal with the conceptual confusion between yaw angle (_) and sideslip angle (_). Seldom is the magnitude of the resulting _ equal to a sudden _ during a yaw upset. The cross effects of a sideslip angle result in a yawing moment and a rolling moment. If the aircraft experienced a rolling motion tightly coupled with the sudden yaw upset, then the Euler yaw angle (the rotation angle of the aircraft about its vertical axis) would be greater than the indicated FDR Heading change.

In their analysis of the AA Flight One accident, the CAB summarized some of the effects of sideslip:

It must also be borne in mind that swept-wing airplanes are subject to a more pronounced roll - yaw coupling than straight-wing aircraft. This roll due to yaw was referred to as “dihedral effect” on straight wing airplanes. When a swept-wing airplane with dihedral yaws, not only is the advancing wing at a higher angle of attack but it also presents a greater span to the airstream. Also, the retreating wing is less effective due to the change in airflow to a more spanwise direction. The lift differential developed by the swept wings is therefore higher and produces a greater rolling moment than would by experienced with a straight-wing airplane under similar conditions. It follows therefore that the roll due to yaw input of the rudder is much more pronounced on swept-wing than on straight-wing aircraft. [57]

The above excerpt mentioned the difference in the angle of attack experienced by each wing when sideslip affects an aircraft with dihedral. The Board’s analysis of the sideslip aspects of the TWA 841 upset (on AAR page 31, second paragraph) did not address this differential of angle of attack.

In a discussion of sideslip, the Board should painstakingly avoid misuse of terms describing the direction of sideslip. Depending upon its direction, the consequence of a sideslip angle can be stabilizing or de-stabilizing. Generally engineering textbooks describe the direction of the sideslip angle as positive or negative. In the report on the AA Flight One accident, the CAB used the nomenclature nose right sideslip or nose left sideslip. Pilots often use the terms left sideslip or right sideslip. With such a variety of terms used by various segments of aviation, confusion can easily develop when discussing the direction of sideslip. The following excerpt is from the DOT accident investigation textbook:

To understand the difference between yaw and slip, visualize an aircraft in which full right rudder is applied; the aircraft yaws to the right, but it slips to the left, regardless of its attitude in unstalled flight. [58]

The cross effects of sideslipmay be undesirable. The following passage from the USAF’s Aerodynamics for Pilots addresses a concept applicable to the TWA 841 upset. It also shows the terminology used by pilots in discussions of the direction of sideslip.

Proverse rollrefers to the roll encountered when an aircraft is yawed. In this case, an aircraft is put in a right yaw by applying right rudder. This creates aleft sideslip angle. Recalling that a negative sideslip angle produces a positive rolling moment, the airplane rolls to the right. The wing does not know whether it is level or not. It always responds to a negative sideslip with a positive roll. Another factor that contributes to the proverse roll is the difference in the velocities of each wing. In the above case, the left wing has a greater velocity than the right wing because of the yawing motion about the airplane center of gravity. This increased velocity increases the lift force on the left wing, which causes a positive roll only as long as the aircraft is yawing. [59]

The NTSB consistently used the term “right sideslip” (on pages 30, 31, 32 and 34, of the AAR) in their report on the TWA 841 upset. The Board used that term to describe the condition of the accident aircraft as it experienced a negative sideslip angle. As is shown in the preceding excerpts, segments of the aviation community perceive the NTSB’s usage of the term “right sideslip” to be incongruous. The Board presented flight test data regarding B727 sideslip in one paragraph on AAR page 31; it was unclear to the reader if the resulting sideslip angle produced de-stabilizing effects. In a revised sideslip analysis of the TWA 841 upset the Board should clearly describe for the reader the direction of sideslip and any de-stabilizing effect of sideslip.

Section 3.CONCLUSIONS

The accident aircraft was not controllable for an interval of approximately one minute. From the beginning of the upset at FL390 until the pilots extended the landing gear, the aircraft was in a sustained nose right sideslip, yawing and rolling to the right.

Full deflection of the cockpit controls, directional and lateral, did not stop the persistent yawing and rolling motion of the aircraft.

After an interval of about one minute of an uncontrollable right rolling motion, while diving through approximately 12500 feet, the pilot commanded “Gear Down.” During the extension swing the right main gear was subjected to extreme side forces. The side loads on the swinging right gear caused structural damage to components of that gear, and the adjacent inboard flap. This damage caused a major hydraulic line to rupture. The rupture of that hydraulic line resulted in the sudden loss of the Hydraulic System “A”.

It was this fortuitous failure that saved the aircraft and passengers. The control problems of TWA 841 very nearly ended with the same fatal consequences as AA Flight One in 1962. The convenient position of the moon was another fortuitous factor that contributed to the successful recovery from the dive. Though intermittently obscured by clouds, the quarter moon was 48 degrees above the southwest horizon. During the pull-out from the dive the pilot used the moon as his main attitude reference. The aircraft’s artificial horizons showed only the black (near vertical nose down) indication, and were of little usefulness during the recovery pullout.

Due to its limited structural strength, an extended #7 slat could not have sustained the loads experienced during the prolonged interval of uncontrollable flight. Hydraulic pressure held the #7 slat in the more secure retracted position until Hydraulic System “A” failed, during the extension of the right landing gear. The #7 slat then immediately extended and separated. The crew sensed the sudden extension and structural failure of the #7 slat as an element of the explosion heard after they extended the gear.

The Board should have commended the pilots of TWA 841 for their persistence. The pilots persevered in their efforts to gain control of the aircraft, trying even the unusual act of extending the landing gear during the uncontrollable dive. This act of last resort, combined with the good fortune of the resulting hydraulic failure, permitted the pilots to regain control of the aircraft.

The Board refused to acknowledge the success of the pilots. In its final Report of June 1981, the Board accused the pilots of illegal and irrational activities. The Board accused the flight crew of first causing the upset, and then making “untimely use of the flight controls.” These actions by the Board may have served to protect the major U.S. aircraft manufacturer.

The NTSB made numerous errors during its investigation of the TWA 841 accident.

The NTSB failed to utilize the wreckage distribution chart to re-create the sequence of failures.

A factor that hampered the careful cataloging of evidence was the early assumption that an extended #7 slat had caused the loss of control. The NTSB’sAircraft Accident Report reflected investigative mistakes previously mentioned. Damaged components such as the separated spoiler panel and flap track “canoe” fairing were mis-identified. The airline had removed rudder actuators and yaw damper transfer valves from the accident aircraft. The NTSB failed to impound or examine those parts.

The Safety Board listed twenty-two “Findings” in the section 3.1 of their Report. The direct evidence did not support several of the “Findings.” The Board should delete“Findings” numbered 10, 12, 13, 14, 15, 16, 18, and 19. Those “Findings” were erroneous conclusions based upon the incorrect assumption that an extended slat had caused the upset. For the same reasons the Board should discard section 3.2 of their Report regarding “probable cause.”

In the “Conclusions” section of the NTSB Report, Finding #6 erroneously stated that there was no irregularity of any flight control system that might have caused a lateral control problem. Post accident “functional checks” of flight controls failed to identify any faults. The NTSB’s Report (page 3) mentioned that the crew had observed the Lower Yaw Damper Flag. The Board ignored that important evidence. Investigators never inspected possible sources of signal distortion due to defective connectors. The Board failed to perform the proper tear-down analysis of the two main rudder actuators. Similarly, the Board failed to perform the proper tear-down analysis of the slat retract locking mechanism for the #7 slat actuator, or of components of the CVR.

The Board seemed completely unaware of the lessons learned from the investigation of the AA Flight One accident. [60]The investigators should have been skeptical about the thoroughness of ground functional tests. Most components that experience an intermittent malfunction can pass the functional test. It is the proper role of an accident investigator to accomplish the appropriate tear down analysis. The NTSB accident investigators should have completed that tear down analysis—independent of the manufacturer. The TWA 841 accident involved loss of directional control. The NTSB investigators should have accomplished the tear down analysis of two rudder actuators [61]removed from the accident aircraft. The NTSB investigators failed to accomplish the appropriate tear down analysis.

ERASURE OF THE COCKPIT VOICE RECORDER

The Board was quick to assume that the CVR had been bulk erased. On page 33 of the AAR, the NTSB stated that the Captain erased the CVR. The Board should delete that accusation from the NTSB Report. The two other pilots stated that they did not see the captain press the CVR Erase Switch. The Captain stated that he did not recall such an action. It is likely that NOBODY pressed the CVR Erase Switch. Due to the hazardous nature of the emergency landing, with reports of sparks and leaking fuel, [62]the pilots were immediately forced to coordinate between the rescue personnel and the flight attendants to start the passenger egress. Then as the pilots completed their post-flight duties their conversation about the upset [63]was recorded by the CVR as it initiated recording for a nine minute interval. Contrary to the NTSB’s insinuations summarized on AAR page 33, the pilots always cooperated with authorities and tried to provide them with information about the upset. The Captain willingly spoke with two FAA inspectors in his hotel room in the middle of the night, within hours of the accident. One FAA inspector quoted the Captain:

You know, we have always been told not to talk to you (FAA) when something happens like this (the N840TW accident), but I can’t see it, as I don’t have anything to hide. I really don’t know what happened. [64]

The Safety Board members had no recent experience with faults in passive, electro-mechanical aircraft systems. The NTSB investigators failed to investigate the probable fault that triggered the bulk eraser. That fault, which energized the bulk erase coil, probably occurred soon after landing with the change-over of electrical power to the APU generator. The NTSB failed to accomplish the appropriate tear down analysis of the voice recorder system. Instead, Board members accused the crew of erasing the CVR. The Board combined this innuendo concerning the CVR erasure, along with its assumption that the crew had attempted to extend flaps during cruise, to build a case in which the Board attributed the cause of the accident to crew actions. This combination of disparaging presumptions was the origin of a bias planted in the minds of the investigators; this bias fouled each phase of the NTSB’s investigation.

NTSB Report Section 3.2,PROBABLE CAUSE

The Board should study the CAB’s conclusions regarding the cause of the AA Flight One accident. The Board should study the causes of the MAC 59402 mishap of November 1976.

The probable cause of the TWA 841 accident was a complex interaction that involved the tightly coupled response of lateral and directional flight controls on the B727-100 aircraft. The failure of a component in the outboard right aileron induced roll and yaw. A malfunction of the automatic rudder control system produced further yaw, sideslip and roll. This interaction lead to a loss of control from which recovery was effected only after hydraulic power to the lower rudder failed. Contributing to the seriousness of this complex interaction was the high cruise altitude and night time conditions. Prior to this accident, due to the rising cost of fuel, the airline had encouraged fuel conservation procedures involving high cruise altitudes.

Section 4,RECOMMENDATION

The Board should discuss the need to inform B727 crews about the risks of encountering common mode failures during high altitude flight. The autopilot and the yaw dampers can induce upsets or impair recovery when at high altitude. The ability of the pilot to first identify and then correct such an upset diminishes when visual clues are lacking, especially at night or under conditions of marginal visibility. Due to the lack of fault annunciation for the yaw dampers and poor instrumentation of the rudder control system, it is unlikely that a pilot could quickly identify the source of a sudden yaw induced rolling motion. A B-727 anomaly known as control wheel ratcheting might restrict the lateral control capability available to the pilot during such an upset. The resulting un-centered neutral point for the control wheel might contribute to the to the instability when the pilot attempts to center the controls. Utilizing the highest possible cruise altitude for the B727-100, as a fuel conservation procedure, may be contrary to safety.

Some members of the Board may contend that investigative errors that occurred a decade ago are now unimportant. Each member of the Board should carefully consider the nature of this case — investigators ignored direct evidence. Instead, the Board utilized circumstantial evidence created at the manufacturer’s facilities after the accident. The NTSB must acknowledge numerous investigative errors.

NTSB RECORD KEEPING

Throughout the investigation, the NTSB seemed to be completely unaware of related mishaps. The NTSB had analyzed the FDR from a similar high altitude upset, MAC 59402, in 1976. [65]Yet the NTSB failed to identify the similarity of the upsets. The NTSB failed to associate the TWA 841 upset with numerous other, less spectacular, B727 rudder induced incidents. The NTSB seemed unaware that the accident aircraft, N840TW, suffered a prior near loss of control. That incident occurred less than two years prior to the accident. In accordance with NTSB Rules, §830.5(a), the Board required airlines to notify the NTSB Field Office, immediately after an incident involving a malfunction of flight controls. Yet the NTSB showed no record of numerous incidents. In some cases, even after the airline properly notified the NTSB, the NTSB showed no record of the incident. [66]

Lacking the history of such related incidents, the Board failed to prepare its investigators for their investigative responsibility. Many line pilots had experienced minor spiral instability episodes, associated with faulty rudder control systems. However, without proper record keeping, the NTSB showed no knowledge those related incidents. The Board’s negligence caused the investigators to rely upon the manufacturer for a suggestion of probable cause.

The Board should retain a record of all flight control incidents. Accident investigators from various parties depend upon the Board to perform that function. In many cases the precise cause of a flight control incident goes undetermined. It is likely that such an unexplained incident would not be recorded as a Service Difficulty Report. The NTSB should require full reporting of any flight control incident in accordance with the details of NTSB Rule §830.15. When initiating an accident investigation the investigators should check those NTSB records for related incidents. Accident investigators should not have to rely upon the manufacturer for such services.

THE NTSB’S OBLIGATION

The falsely accused TWA crew deserves an apology from the manufacturer and from the NTSB.

NTSB ReportAppendix B erroneously stated the employment date of the first officer as 1969. In fact, the F/O had been employed by the airline since December 1966. [67]

A brief table correlating events of the upset with FDR information is provided on the following pages.

_




“C” Check 9, 38

“LE Flap” annunciation 28


Footnotes

[1]The response of the CVR system to a change of the electrical power source was described by Fairchild’s CVR expert.

[2] This designation for the missing spoiler panel was found to occur elsewhere in the records of the investigation. The TWA Engineering report of Findings and Work Accomplished (prepared by R. J. Sorenson) was included in the docket. Page 2 of that report used the incorrect designation “#10 flight spoiler,” but correctly associated the missing spoiler panel with the right wheel well area. The Structures Group Investigation Factual Report on page 2 used the same incorrect label for that spoiler panel. The Boeing Report correctly referred to the missing spoiler as “#10 spoiler panel” which is an alternative designation for the #6 flight spoiler .

[3]Photographs of the accident aircraft were included in the NTSB's docket and as evidence in the post accident litigation. One photograph from the litigation, Defendant's Exhibit TWA-28, clearly showed the upper surface of the right wing. This photographic evidence was retained by the trial attorney, Mr. Donald Chance Mark, 4200 Multifoods Tower, 33 South Sixth Street, Minneapolis, MN, 55402. Telephone (612) 338-0661.

[4]Erroneous information may have confused the NTSB investigators. The manufacturer had supplied the NTSB with an analysis of a slat locking mechanism. (See the NTSB docket, regarding an analysis completed 28Apr80; submitted to the NTSB under a Boeing cover letter dated 15Dec80.) In that analysis, the manufacturer had analyzed possible faults in a newer, re-designed, retract locking switch. The newer switch was unlike the mechanism (for that slat actuator) on N840TW. Thus the NTSB may have erred when it denied that there were any spring forces acting from the retract locking switch.

[5]Boeing Operations Manual Bulletin 75-7 ; included in NTSB Aircraft Accident Report 81-8, pages 49 and 50, as Appendix E.

[6]CAB, Aircraft Accident Report , American Airlines Flight One, (CAB Report Released January 15, 1963), 30-33 and 51-52. Boeing 707-123B, N7506A, crashed (fatal) into Jamaica Bay, Long Island, New York; on March 1, 1962.

[7]CAB, Aircraft Accident Report , American Airlines Flight One, 33.

[8]CAB, Aircraft Accident Report , American Airlines Flight One, 51-52. [Emphasis added.]

[9] Boeing Document D6-44357, Post Flight Conference Transcript for Test 83-1, 10-2-80, page 83-1-D2.

[10] NTSB, Addendum to Performance Group Report (January 7, 1981), 2.

[11] P. J. Neufeld and N Colman, "When Science Takes the Witness Stand," Scientific American, 262:5 (May 1990), 46-53.

[12]Hogue to Kampschror letter, May 11, 1981; attachment page 2, fourth paragraph. Emphasis added.

[13] NTSB, Structures Group Investigation Factual Report (Washington, D.C.: NTSB Bureau of Technology, May 10, 1979), 1.

[14] The term left sideslip is used here to describe the case when the relative wind is to the left of the nose. This is defined as a negative sideslip angle . This terminology is taught to pilots, and is used in US Air Force ATC Manual 51-3, Aerodynamics for Pilots (Randolph AFB, TX, 1 July 1970), 127. However, following the convention used by the CAB in the AA Flight One AAR, the term nose right sideslip will be used herein to describe a negative sideslip angle . The term proverse roll is also defined in the above reference.

[15] TWA’s B727 FHB; page 14.01.03, dated October 1, 1989.

[16]TWA’s B727 FHB; page 9.01.09, dated January 1, 1983.

[17] TWA’s B727 FHB; page FPS 20.24, dated October 15, 1986.

[18] Boeing Document D6-8095, Revision 9-30-63, page 4.3-2.

[19]Charles Perrow, NORMAL ACCIDENTS: Living with High-Risk Technologies (New York: Basic Books, Inc.,.1984), 217.

[20] MM 22-00, page 1.

[21] MM 22-00, page 10.

[22] MM 22-00, page 16.

[23]DOT Transportation Safety Institute, Aircraft Accident Investigation Procedures and Techniques Textbook, (Oklahoma City, OK: FAA Aeronautical Center, May 1978), 172. [Emphasis added.]

[24]Figure 4, on page A-10 of the Boeing Report, erroneously labeled the separated canoe fairing as the “#6 Flap Track Fairing.” That mistaken designation suggested that the separated fairing had ripped from the outboard track of the inboard flap. Photographic evidence showed clearly that it was the mid section of the #5 flap track fairing which was missing from the accident aircraft.

[25] Some incidents of slat separation, at speeds between 310 and 360 KIAS, were recorded as SDR’s: EAL/N8139 on 1May69, NWA/N499 on 2Sep69, NAL/N4746 on 10Dec69, AAL/N1977 on 18Jul72, AAL/N6808 on 30Aug72, UAL/N7088 on 12Dec72, and EAL/N81590 on 26Dec72.

[26] See USAF Mishap Report, Number 761111201. Mishap involved a Lockheed C-141A, MAC 59402, on 11 NOV 1976, near the Campbell Islands, British Columbia, Canada. The mishap report is available from HQ AFISC (Inspection and Safety Center); Norton AFB, CA. The report described an upset that began at 41000 feet. The period of instability included yaw, roll-over, and a near vertical dive. All control inputs were ineffective until approximately 24000 feet.

[27] NTSB Deposition of [the Flight Engineer] (Inglewood, Ca, April 12, 1979); Transcript Page 16. [Emphasis added.]

[28] Maintenance Records Group, Group Chairman’s Factual Report of Investigation (part of the NTSB docket, report dated May 4, 1979), page 3, lines 1 through 5.

[29] Ibid., attached Non-Routine Repair Record, first page, photocopy of Card No.15, Area 600, date Feb 29 [sic].

[30] Systems Group, Systems Group Chairman Report of Investigation (part of the NTSB docket, report dated May 7, 1979), page 2, last paragraph.

[31]Excerpt from Remarks section of an Incident Report, attached to FAA Form 2810, by H. P. Gassaway, of office SO-GADO-6. [Emphasis added.]

[32] Photographic evidence shows that it was the I/B flap I/B track, the #5 flap track fairing, that separated from the aircraft.

[33] Boeing 727 Maintenance Manual , Section 27-00 page 1202.

[34]NTSB Deposition of H. G. Gibson, (Inglewood, CA, April 12, 1979); page 26.

[35]NTSB Deposition of H. G. Gibson, (Inglewood, CA, April 12, 1979); page 27.

[36]NTSB Deposition of H. G. Gibson, (Inglewood, CA, April 12, 1979); pages 27 and 28.

[37]Taken from a description of spiral divergence; Francis J. Hale, Aircraft Performance, Selection and Design [New York: John Wiley & Sons], 263.

[38]NTSB Docket, 727 Wing Buffet and Pressure Distribution , page 2, first and second paragraphs. Cover letter dated April 16, 1980; from Boeing's H. P. Hogue to NTSB's L. D. Kampschror.

[39]NTSB Deposition of H. G. Gibson, (Inglewood, CA, April 12, 1979); page 27.

[40] NTSB, Human Factors Specialist's Report of Investigation (June 22, 1979), page 3.

[41]NTSB, Human Factors Report , page 4.

[42] NTSB, Human Factors Report , Attachment I-1.

[43]NTSB Deposition of H. G. Gibson ; page 38.

[44]NTSB Deposition of H. G. Gibson ; page 42.

[45]NTSB, Human Factors Specialist's Report of Investigation (June 22, 1979), Attachment I-2.

[46]NTSB, Addendum to Performance Group Chairman’s Report of Investigation (May 27, 1981), page 3, and Attachment 1.

[47]Boeing 727-31 Maintenance Manual , 22-00 page 45

[48]NTSB Aircraft Accident Report 81-8, pages 49 and 50, included as Appendix E.

[49]NTSBAircraft Accident Report 81-8, page 24: “. . . the flight crew . . . saw no lights in the cockpit that indicated an unlocked leading edge device . . .”

[50]NTSB AAR, page 24.

[51] NTSB, Response to Petition For Reconsideration , dated Dec 15, 1983; re ALPA petition on NTSB-AAR-81-8.

[52]NTSBDeposition of H. G. Gibson, (Inglewood, CA, April 12, 1979); page 29.

[53]Boeing Commercial Airplane Company, Postflight Conference, Test 83-1 , transcript (Seattle, WA: verbatim reporter Gus Baisch, October 2, 1980) D6-44357, page 83-1-D4.

[54]NTSBDeposition of H. G. Gibson, (Inglewood, CA, April 12, 1979); page 27.

[55] In a Boeing flight test of 1975 with an extended #2 slat, the intensity of the associated buffet had caused the test crew to abandon their testing while climbing through 33400 feet at Mach 0.81. The test log recorded that the buffet intensity had caused “extreme cowardice,” preventing the test crew from further testing of that configuration.

[56]NTSBAircraft Accident Report 81-8, page 31.

[57]CAB,Aircraft Accident Report ,American Airlines Flight One, (CAB Report Released January 15, 1963), 35-36.

[58]DOT Transportation Safety Institute, Aircraft Accident Investigation Procedures and Techniques Textbook,(Oklahoma City, OK: FAA Aeronautical Center, May 1978), 131.

[59]US Air Force ATC Manual 51-3, Aerodynamics for Pilots (Randolph AFB, TX, 1 July 1970), 127.

[60]CAB,Aircraft Accident Report ,American Airlines Flight One, (CAB Report Released January 15, 1963), 22-30.

[61]DOT Transportation Safety Institute, Aircraft Accident Investigation Procedures and Techniques Textbook,(Oklahoma City, OK: FAA Aeronautical Center, May 1978), 134.

[62]NTSB,Deposition of H. G. Gibson, (Inglewood, CA, April 12, 1979); page 57. NTSB, Deposition of the Flight Engineer , pages 20-21. NTSB, Deposition of the First Officer, page 19-20.

[63]NTSB, Cockpit Voice Recorder Transcript (Washington, D.C., April 16,1979), 4-8.

[64]NTSB docket, excerpt, FAA Form 3112 report, by R.L. Montgomery, FAA Airworthiness Inspector, at Air Carrier District Office 35.

[65] NTSB Flight Data Recorder readout, Project/Report No. 77-7, readout date 11-19-76.

[66]The airline reported a high altitude upset, that occurred near Dayton, after midnight on December 17,1989; aboard TWA1070. The NTSB's Office of Public Inquiries showed no record of the incident.

[67] NTSB, Operations Group Report, page 4.


INDEX

A , B , C , D , E , F , G , H , I , L , M , N , O , P , R , S , T , V , W , Y


autopilot disconnect function, 1
MAC 59402 mishap, 1
TWA B727 Flight Handbook , 1
A
AA Flight One accident, 1, 2, 3
AA Flight One accident, 1
access panel damage, 1
actuator piston, 1
actuator rod, 1, 2
actuator rod ruptured, 1
Aerodynamics for Pilots , 1
aileron damage, 1, 2
aileron hinge mid rib bolt, 1
aileron-spoiler mixer, 1, 2
air data sensor, 1
air loads, 1
air loads on the slat, 1
aircraft damage illustration, 1
aircraft parts, 1
aircraft’s flight path, 1
airspeed decrease, 1
airspeed trace comparison, 1
angle of attack , 1
annunciation, 1, 2, 3, 4, 5, 6
annunciation , 1
annunciation for yaw damper faults, 1
autopilot, 1, 2
autopilot false disconnect, 1
autopilot false disconnect , 1
autopilot heading hold function, 1
autopilot lateral inputs, 1
autopilot’s manual mode, 1
Autopilot,, 1
B
B-727-200, 1
bias, 1, 2
body length, 1
Boeing OMB 75-7, 1
Boeing Operations Manual Bulletin 75-7, 1, 2
Boeing Ops Manual Bulletin, 1
buzz, 1
C
C Check, 1, 2
Capt. McIntyre, 1
change-over of electrical power, 1
circumstantial evidence, 1, 2
Cockpit Voice Recorder, 1
common mode failures, 1
Component Failure Diagram, 1
control wheel, 1
control wheel ratcheting, 1, 2, 3
coupled interaction of yaw and roll, 1
cross effects of .i.sideslip, 1
cross effects of a sideslip angle, 1
CVR erasure, 1
D
data plots, 1
debris, 1, 2
differential damage, 1
dihedral effect, 1, 2
direction of sideslip, 1, 2
directional control system, 1
directional gyro, 1, 2
discrepant rudder deflection, 1
distorted bolt holes, 1
divergent, 1
Duane Yorke, 1
dynamic instability, 1
E
electrical connections, 1
electrical power change-over, 1
electro-hydraulic servo valve , 1
engineered safety feature, 1
engineered safety features, 1
Euler yaw angle, 1
explosion, 1
Extend Stop, 1
Extend Stops, 1
extended . 7 slat
extension of the slat, 1
F
Failure Sequence, 1, 2
Failure Sequence Chart , .i.vibration , 1
fatigue progression, 1, 2
fatigue type fracture., 1
fault, 1
fault analysis of the retract locking mechanism, 1
FDR, 1, 2, 3
FDR Airspeed Trace, 1
FDR G-trace, 1
FDR Heading changes, 1
FDR Heading stylus, 1, 2
FDR heading trace, 1, 2
FDR speed verses slat strength, 1
flap track canoe fairing, 1, 2, 3, 4, 5, 6, 7, 8, 9
flaps, 1
Flight 1070, 1
Flight 5562,, 1
flight controls, 1
flight controls illustration, 1
Flight Data Recorder, 1
Flight Data Recorder (FDR) parameters, 1
flight test, 1
Flutter, 1, 2
flutter margin, 1
fraction of critical damping, 1
fracture of the hinge bolt, 1
frequency response, 1
FTI, 1
functional checks, 1
functional tests, 1
G
G forces, 1
G-trace, 1, 2
g-trace (vertical acceleration) frequency comparison, 1
gear unsafe indications, 1
gimbal assembly, 1
gimbal error, 1
gyro gimbal error, 1, 2
H
heading gyro tests, 1
heading trace, 1, 2, 3, 4
hydraulic leakage, 1, 2, 3, 4, 5
hydraulic leakage , 1
hydraulic line damage, 1
hydraulic line rupture, 1, 2, 3
hydraulic line rupture ), 1
Hydraulic pressure, 1, 2
Hydraulic System, 1
I
incident on December 16, 1989, near Dayton, Ohio, 1
incidents, 1, 2
L
L. D. Kampschror, 1
Landing Gear, 1
Landing Gear Damage, 1, 2, 3, 4
landing gear side brace, 1
lateral control deflections, 1
lateral dynamic instability, 1
lateral fracture direction, 1
LEFlap annunciation, 1
light control incident , 1
load bearing strength of an extended slat, 1
lower rudder actuator, 1, 2
Lower Yaw Damper Flag, 1
M
MAC 59402, 1
Maintenance Records, 1
Maintenance Records Group, 1, 2
metallurgist, 1
moon, 1, 2
N
noise level, 1
nose gear, 1
nose right sideslip, 1
O
outboard aileron, 1
outboard aileron hinge bolt failure, 1
P
parameter relationships, 1
possible gimbal errors in the rate gyros, 1
Post Accident Inspection, 1
Project Race, 1
Proverse roll, 1
R
ratcheting, 1
ratcheting of the aileron control wheel, 1
recorder mechanisms, 1
recorder stylus shifts, 1
recovery profile view (FDR), 1
response to a petition, 1
Retract Lock Piston Spring, 1
Retract Lock Switch, 1
Retract Position Indicating Switch, 1
retracted . 7 slat
right main landing gear, 1
right outboard aileron, 1, 2
rolling moment, 1, 2
Ronson Corporation, 1
rudder, 1, 2, 3
rudder actuator leakage, 1
rudder actuators and .i.yaw damper transfer valves, 1
rudder actuators, and both .i.yaw damper transfer valves,, 1
Rudder and Elevator Position Indicator, 1
Rudder Boost Packages, 1
Rudder control system malfunction, 1
rudder control systems, 1
rudder deflection, 1, 2, 3
rudder displacement, 1
rudder hard-over indication, 1
Rudder Position Indicator illustration, 1
rudder position indicators, 1
Rudder Yaw Damper Engage switches, 1
rudder-fin illustration, 1
rudders, 1
S
seal at the retract switch Sensing Pin, 1
self test features of the CVR, 1
Separated Parts, 1, 2, 3
separation of a slat, 1
Separation of parts, 1, 2
separation of the . 6 flight spoiler
separation of the slat, 1
Service Difficulty Report, 1
Sideslip, 1, 2, 3, 4, 5, 6
sideslip , 1
sideslip angle, 1
sideslip angles of 4.5_ to 6.5_, 1
sideslip condition, 1
sideslip illustration, 1
sideslip induced rolling motion, 1
sideslip trace, 1
signal distortion in defective connectors, 1
simulator testing, 1
Simulator Tests, 1
skydrol , 1
skydrol bathing, 1, 2, 3
skydrol damage, 1
skydrol leakage, 1
slat, 1, 2, 3
slat . 7
slat . 7
slat . 7 pre-existing damage
slat . 7 .i.hydraulic line rupture )
slat . 7
slat . 7
slat actuator, 1, 2
slat actuator mechanical retract lock, 1
Slat Actuator Retract Lock Indicating Switch, 1
slat actuator Retract Lock Keys, 1
slat annunciator, 1
slat box buckled, 1
Slat Extend Position Indicating Switch, 1
slat extended, 1
Slat load bearing strength, 1
slat misalignment, 1
Slat Retract Lock Indicating Switch, 1
slat separation, 1, 2, 3, 4
slat track, 1, 2
Slat Track Extend Stop, 1
slat tracks, 1
slat warning lights, 1
slats, 1
speed brakes, 1
speedbrakes, 1, 2
spiral dive, 1
Spoiler, 1
spoiler blow down, 1, 2
spoiler buzz, 1, 2, 3
spoiler damage, 1
spoiler float, 1
spoiler panel damage illustration, 1
spoilers illustration, 1
stability characteristics, 1
structural components of the slat, 1
stylus anomalies,, 1
sweepback effect, 1
T
T-bolt, 1
tear down analysis, 1
tear-down analysis, 1
trail of debris, 1, 2, 3
Trajectory Analysis, 1, 2
transfer valve, 1
TWA B727 Flight Handbook , 1
TWA flight tests of 1977, 1
V
vertical fin, 1
vertical stabilizer, 1
vibration, 1, 2, 3, 4, 5, 6, 7
vibration , 1, 2
vibration , 1
vibration intensity, 1
W
winds, 1
wing surface, 1
wreckage distribution chart, 1
Y
yaw calibrator, 1
yaw damper, 1, 2, 3
yaw damper coupler, 1
yaw damper couplers, 1, 2, 3
Yaw Damper Fail Flag, 1, 2
Yaw Damper Fail Flag , 1
yaw damper fail flags, 1
yaw damper hardovers, 1
yaw damper inputs, 1
yaw damper signals, 1
yaw damper transfer valves, 1, 2
yaw damper transfer valves,, 1
yaw dampers, 1, 2
yaw induced rolling motion, 1
yaw motions, 1
yaw rate gyro, 1, 2
yaw servo-amplifier, 1
yaw synchronizer, 1
yaw-rate, 1
yaw-rate gyro, 1
yawing moment, 1
yawing moment persisted, 1