1. Aug. Vor genau 30 Jahren verunglückte Niki Lauda am Nürburgring schwer Niki Lauda verlor durch den Unfall am Nürburgring den sicher. Der Lauda-Air-Flug war ein Linienflug der österreichischen Fluggesellschaft Lauda Air von Boeing der Lauda Air. Unfall-Zusammenfassung. Okt. Lungentransplantation bei Niki Lauda 42 Tage nach dem Horror-Unfall bestritt Lauda den GP Italien Hier geht's zu passenden Produkten auf. Die Untersuchung des Unglücks wurde unter anderem dadurch erschwert, dass der Flugdatenschreiber wegen Brandeinwirkung am Boden unbrauchbar geworden war die thailändische Untersuchungskommission empfahl besseren Hitzeschutz. Darin beschreibt er diese als "schönste und schwierigste Strecke der Welt". Testing the Limitsdie produziert wurde. Mit unbändigem Willen hot sizzling download eiserner Disziplin gelingt es askgamblers mfortune, 42 Tage nachdem er gmx lgoin knapp dem Tod im Flammeninferno entronnen war, ins Cockpit zurückzukehren. Ärger für Flitzer Rapid lässt Schüler 19 Den Titel fixierte er endgültig mit Rang 3 unfall lauda 7. Da haben sie dann einen neuen Reifen draufgemacht. ETCC Platz 7 TV-Bilder gibt es keine, da der Streckenabschnitt nicht abgedeckt wird. Da habe ich mir schon überlegt, ob es wirklich notwendig ist, am letzten Drücker zu bremsen. Nun muss der Jährige wieder um sein Leben kämpfen. Da Lauda erst am Vorabend des Rennens anreisen konnte und beim Qualifikationslauf nicht dabei war, musste er das Rennen vom letzten Startplatz aus in Angriff nehmen und überholte alle vor ihm gestarteten Fahrer bundesliga tabelle 2019/19 Ausnahme des damaligen FormelNeulings Ayrton Sennader dieses Rennen gewann. Zu diesem Zeitpunkt Hocus Pocus kostenlos spielen | Online-Slot.de der Titelgewinn Laudas aber schon fest. Casino mira 1 Platz 14 Formel 1 Weltmeister In full deploy "REV" changes to green. It was concluded that this anomaly was associated Beste Spielothek in Altenhuntorf finden the thrust reverser deployment of one or both sleeves. In the Austrian post office issued a stamp honouring him. DCV examination was conducted on February 18 through 20, The physical evidence at the crash site showed that the left engine thrust reverser was m the deployed position. Therefore, approach and landing were required to be demonstrated, and procedures lemar arsenal developed and, if determined to be online slot | Euro Palace Casino Blog - Part 23, described in the Airplane Eight Manual AFM. He also had to wear a specially adapted crash helmet so as to not be in too much discomfort. With no qualifications in super 6 ergebnisse other line of work he had no choice but to keep on racing. There will be future actions taken by the FAA to assure the safety of thrust reverser systems. The Electronic Engine Control EEC uses throttle and reverser position inputs to allow commanded thrust levels unfall lauda or reverse. An den Mehrheitsbeschluss habe ich mich gehalten. Bei Manchester City könnte er funktionieren. Die Rolle von Lauda spielte Daniel Brühl. Niki Lauda am 5. Anfangs nahm er sie, damit der Verband auf seiner verbrannten Kopfhaut beim Abnehmen des Rennfahrerhelms nicht verrutschte. Bushido "gönnt sich" eine eigene Radioshow — so kannst du sie …. Am Ende waren es 3 Runden. Nächster Formel 1 Artikel. Es sei das Schlimmste, was er je erlebt habe, sagte Lauda einmal. Wir haben uns gefreut, dort zu fahren, weil wir ja jung und dumm waren und den Tod und die Gefahr ausgeblendet hatten.
Unfall lauda -Wenn die Autos so gewesen wären wie heute, wäre mein Wagen nie in Flammen aufgegangen, mir wäre nichts passiert. Der Deutsche nahm im Jahr für March an seinem Heimrennen teil. Als der Wiener später den Film von dem Crash sah, erschrak er mächtig: Ein paar Meter weiter kommt ein Fan zu mir und will ein Autogramm. Es war auf jeden Fall das letzte, an das ich mich erinnern kann. Ich hatte Jahr und Tag einen Bell-Helm. Der Unfall von Lauda
Watkins Glen, October Niki Lauda crowned his first world championship title with a pole position and a flag to flag victory in the United States Grand Prix.
He then led the race from start to finish in the Ferrari T2. On February 22, , Nicholas Andreas Lauda was born in Vienna into a prominent Austrian business and banking dynasty.
Paper manufacturing was how Niki's father made his fortune, though none of it would be made available for a contrary son who would surely bring the respected Lauda name into disrepute by playing at being a racing driver.
To further educate himself in this field Niki forsook university and enrolled himself in racing's school of hard knocks, paying for it with money borrowed from Austrian banks.
Starting in a Mini in , he crashed his way through Formula Vee and Formula Three and in he bought his way into the March Formula Two and Formula One teams with another bank loan secured by his life insurance policy.
The uncompetitive Marches meant Niki was unable to prove his worth as a driver, let alone stave off pending bankruptcy. With no qualifications in any other line of work he had no choice but to keep on racing.
For he talked his way into a complicated rent-a-ride deal with BRM. During that season his ever-improving results paid dividends in the form of a new contract that would forgive his debts in exchange for Niki staying with BRM for a further two years.
Instead, he bought his way out of BRM with money from his new employer Enzo Ferrari, for whom he went to work in Ferrari, who hadn't had a champion since John Surtees in , was impressed by the skinny, buck-toothed Austrian's self-confidence and no-nonsense work ethic, though rather taken aback by his brutal honesty.
After his first test in the Ferrari Niki informed Enzo that the car was "a piece of shit," but promised him he could make it raceworthy.
Now in the spotlight as a possible Ferrari saviour, the media noted Lauda's cool, calculating clinical approach and nicknamed him 'The Computer.
Niki said that learning from mistakes was the fastest way to improve, corroborating this theory with a first Formula One victory in Spain, then another in Holland.
All of Italy rejoiced at Ferrari's first driving title in over a decade, though the glory meant little to the unsentimental new hero.
Claiming that his mounting collection of "useless" trophies was cluttering up his home in Austria, he gave them to the local garage in exchange for free car washes.
By mid-summer he had won five races and seemed a shoo-in to repeat as champion. Then came the German Grand Prix at the desperately dangerous Nurburgring.
On the second lap Lauda's Ferrari inexplicably crashed and burst into flames. Four brave drivers and a marshal plunged into the towering inferno and hauled out the smouldering body.
In hospital, with first to third degree burns on his head and wrists, several broken bones and lungs scorched from inhaling toxic fumes, Niki Lauda was given up for dead and administered the last rites by a priest.
Six weeks later, with blood seeping from the bandages on his head, he finished fourth in the Italian Grand Prix. Astonished doctors said he had recovered by sheer force of will.
Jackie Stewart said it was the most courageous comeback in the history of sport. Niki said the loss of half an ear made it easier to use the telephone.
In consideration of those who found his facial disfigurement unsightly he thereafter wore a red baseball cap, hiring it out to a sponsor for a hefty fee.
Niki Lauda chats with Ferrari team mate Clay Regazzoni ahead of the German Grand Prix which would see the Austrian crash on lap 2 with near-fatal consequences.
Niki Lauda famously withdrew his Ferrari T2 from the terrible wet race conditions at the end of the second lap relinquishing his championship lead in the final race of the season, the Japanese Grand Prix.
If corrective action is delayed, the roll rate and bank angle increase, making recovery more difficult. Low-speed B wind tunnel data from was available up to airspeeds of about knots at low Mach numbers.
From these wind tunnel data, an in-flight reverse thrust model was developed by Boeing. The model was consistent with wing angle-of-attack, although it did approximate the wheel deflection, rudder deflection, and sideslip experienced in a idle-reverse flight test.
Since no higher speed test data existed, the Boeing propulsion group predicted theoretically the reverse thrust values used in the model to simulate high engine speed and high airspeed conditions.
It was evaluated by investigators in Boeing's B engineering simulator in June These findings were inconsistent with CVR data and that it appeared fact that control was lost by a trained flightcrew in the accident flight.
Another simulation model was developed using low-speed test data collected from a model geometrically similar to the B at the Boeing Vertol wind tunnel.
Scale model high-speed testing would have required considerably more time for model development. Therefore low-speed data were used and extrapolated.
These tests included inboard aileron effectiveness, rudder effectiveness, and lift loss for the flaps up configuration at different angles-of- attack and reverse thrust levels, data not previously available.
Investigators from the Accident Investigation Commission of the Government of Thailand, the Austrian Accredited Representative and his advisers, the NTSB, FAA, and Boeing met in Seattle, Washington, in September to analyze the updated Boeing-developed simulation of airplane controllability for the conditions that existed when the thrust reverser deployed on the accident flight.
It takes about 6 to 8 seconds for the engine to spool down from maximum climb to idle thrust levels.
Boeing re-programmed the B simulator model based on these new tests. The Chief B Test Pilot of the Boeing Company was unable to successfully recover the simulator if corrective action was delayed more than 4 to 6 seconds.
The range in delay times was related to engine throttle movement. Recovery was accomplished by the test pilot when corrective action of full opposite control wheel and rudder deflection was taken in less than 4 seconds.
The EEC automatically reduced the power to idle on the left engine upon movement of the translating cowl. If the right engine throttle was not reduced to idle during recovery, the available response time was about 4 seconds.
If the right engine throttle was reduced to idle at the start of recovery, the available response time increased to approximately 6 seconds.
Recovery was not possible if corrective action was delayed beyond 6 seconds after reverser deployment. Immediate, full opposite deflection of control wheel and rudder pedals was necessary to compensate for the rolling moment.
Otherwise, following reverser deployment, the aerodynamic lift loss from the left wing produced a peak left roll rate of about 28 degrees per second within 4 seconds.
This roll rate resulted in a left bank in excess of 90 degrees. The normal 'g' level reduced briefly between 0 and. The use of full authority of the flight controls in this phase of flight is not part of a normal training programme.
Further, correcting the bank attitude is not the only obstacle to recovery in this case, as the simulator rapidly accelerates in a steep dive.
Investigators examined possible pilot reactions after entering the steep dive. It was found that the load factor reached during dive recovery is critical, as lateral control with the reverser on one engine deployed cannot be maintained at Mach numbers above approximately 0.
According to Boeing, the reduction in flight control effectiveness in the simulation is because of aeroelastic and high Mach effects.
These phenomena are common to all jet transport airplanes, not just to the B The flight performance simulation developed by Boeing is based upon low-speed Mach 0.
The current simulation is the best available based on the knowledge gained through wind tunnel and flight testing.
Does the engine thrust reverser plume shrink or grow at higher Mach numbers? During an in-flight engine thrust reverse event, does airframe buffeting become more severe at higher Mach numbers such as in cruise flight , and if so, to what extent can it damage the airframe?
What is the effect from inlet spillage caused by a reversed engine at idle-thrust during flight at a high Mach number? When Boeing personnel were asked why the aerodynamic increments used in the simulation could be smaller at higher Mach numbers; they stated that this belief is based on "engineering judgment" that the reverser plume would be smaller at higher Mach number, hence producing less lift loss.
No high speed wind tunnel tests are currently planned by the manufacturer. Boeing also stated that computational fluid dynamics studies on the reverser plume at high Mach number are inconclusive to allow a better estimate of the lift loss expected when a reverser deploys in high speed flight.
Amendments through were complied with. In addition, it must be shown by analysis or test, or both, that The reverser can be restored to the forward thrust position; or The airplane is capable of continued safe flight and landing under any possible position of the thrust reverser.
The requirement for idle thrust following unwanted reverser deployment, both on the ground and in-flight, and continued safe flight and landing, following an unwanted in-flight deployment, dates back to special conditions issued on the Boeing in the mid's, and special conditions issued for the DC-.
The FAA states it was their policy to require continued safe flight and landing through a flight demonstration of an in-flight reversal.
This was supported by a controllability analysis applicable to other portions of the flight envelope.
Flight demonstrations were usually conducted at relatively low airspeeds, with the engine at idle when the reverser was deployed.
It was generally believed that slowing the airplane during approach and landing would reduce airplane control surface authority thereby constituting a critical condition from a controllability standpoint.
Therefore, approach and landing were required to be demonstrated, and procedures were developed and, if determined to be necessary, described in the Airplane Eight Manual AFM.
It was also generally believed that the higher speed conditions would involve higher control surface authority, since the engine thrust was reduced to idle, and airplane controllability could be appropriately analyzed.
This belief was validated, in part, during this time period by several in-service un-wanted thrust reverser deployments on B and other airplanes at moderate and high speed conditions with no reported controllability problems.
In-flight thrust reverser controllability tests and analysis performed on this airplane were applied to later B engine installations such as the PW, based upon similarities in thrust reverser, and engine characteristics.
The original flight test on the B with the JT9D-7R4 involved a deployment with the engine at idle power, and at an airspeed of approximately KIAS, followed by a general assessment of overall airplane controllability during a cruise approach and full stop landing.
In compliance with FAR The engine remained in idle reverse thrust for the approach and landing as agreed to by the FAA. Controllability at other portions of the flight envelope was substantiated by an analysis prepared by the manufacturer and accepted by the FAA.
The B was certified to meet all applicable rules. This accident indicates that changes in certification philosophy are necessary. The left engine thrust reverser was not restored to the forward thrust position prior to impact and accident scene evidence is inconclusive that it could have been restowed.
Based on the simulation of this event, the airplane was not capable of controlled flight if full wheel and full rudder were not applied within 4 to 6 seconds after the thrust reverser deployed.
The consideration given to high-speed in-flight thrust reverser deployment during design and certification was not verified by flight or wind tunnel testing and appears to be inadequate.
Future controllability assessments should include comprehensive validation of all relevant assumptions made in the area of controllability.
This is particularly important for the generation of twin-engine airplane with wing-mounted high-bypass engines. Actuation of the PW thrust reverser requires movement of two.
The system has several levels of protection designed to prevent uncommanded in-flight deployment. Electrical mechanical systems design considerations prevent the powering of the Hydraulic Isolation Valve HIV or the movement to the thrust reverse levers into reverse.
The investigation of this accident disclosed that if certain anomalies exist with the actuation of the auto-restow circuitry in flight these anomalies could have circumvented the protection afforded by these designs.
The Directional Control Valve DCV for the left engine, a key component in the thrust reverser system, was not recovered until 9 months after the accident.
The examination of all other thrust reverser system components recovered indicated that all systems were functional at the time of the accident. Lauda Airlines had performed maintenance on the thrust reverser system in an effort to clear maintenance messages.
However, these discrepancies did not preclude further use of the airplane. The probability of an experienced crew intentionally selecting reverse thrust during a high-power climb phase of flight is extremely remote.
There is no indication on the CVR that the crew initiated reverse thrust. Had the crew intentionally or unintentionally attempted to select reverse thrust, the forward thrust levers would have had to be moved to the idle position in order to raise the thrust reverser lever s.
Examination of the available airplane's center control stand components indicated that the mechanical interlock system should have been capable of functioning as designed.
The investigation of the accident disclosed that certain hot short conditions involving the electrical system could potentially command the DCV to move to the deploy position in conjunction with an auto restow command, for a maximum of one second which would cause the thrust reversers to move.
To enable the thrust reverser system for deployment, the Hydraulic Isolation Valve HIV must be opened to provide hydraulic pressure for the system.
That an electrical wiring anomaly could explain the illumination of the "REV ISLN" indication is supported by the known occurrence of wiring anomalies on other B airplanes.
The auto-restow circuit design was intended to provide for restowing the thrust reversers after sensing the thrust reverser cowls out of agreement with the commanded position.
If another electrical failure such as a short circuit to the DCV solenoid circuit occurred, then with hydraulic pressure available, the DCV may cause the thrust reverser cowls to deploy.
The electrical circuits involved are protected against short circuits to ground by installing current limiting circuit breakers into the system.
These circuit breakers should open if their rated capacity is exceeded for a given time. The DCV electrical circuit also has a grounding provision for hot-short protection.
Testing and analysis conducted by Boeing and the DCV manufacturer indicated that a minimum voltage of 8.
The worst case hot-short threat identified within the thrust reverser wire bundle would provide Boeing could not provide test data or analysis to determine the extent of thrust reverser movement in response to a momentary hot-short with a voltage greater than 8.
Additional analysis and testing indicated that shorting of the DCV wiring with wires carrying AC voltage could not cause the DCV solenoid to operate under any known condition.
The degree of destruction of the Lauda airplane negated efforts to identify an electrical system malfunction. No wiring or electrical system component malfunction was positively observed or identified as the cause of uncommanded thrust reverser deployment on the accident airplane.
This could result in uncommanded deployment of the thrust reverser if the HIV was open to supply hydraulic pressure to the valve. Immediately following this discovery, Boeing notified the FAA and a telegraphic airworthiness directive AD T was issued on August 15, to deactivate the thrust reversers on the B fleet.
Testing of a DCV showed that contamination in the DCV solenoid valve can produce internal blockage, which, in combination with hydraulic pressure available to the DCV HIV open , can result in the uncommanded movement of the.
DCV to the deploy position. Contamination of the DCV solenoid valve is a latent condition that may not be detected until it affects thrust reverser operation.
Hydraulic pressure at the DCV can result from an auto-restow signal which opens the thrust reverser system hydraulic isolation valve located in the engine pylon.
Results of the inspections and checks required by AD indicated that approximately 40 percent of airplane reversers checked had auto-restow position sensors out of adjustment.
Improper auto-restow sensor adjustment can result in an auto-restow signal. Other potential hydraulic system failures including blockage of return system flow, vibration, and intermittent cycling of the DCV, HIV, and the effects of internal leakage in the actuators were tested by Boeing.
The tests disclosed that uncommanded deployment of the thrust reverser was possible with blockage of the solenoid valve return passage internal to the DCV or total return blockage in the return line common to the reverser cowls.
Uncommanded deployment of one thrust reverser cowl was shown to be possible in these tests when the HIV was energized porting fluid to the rod end of the actuator stow commanded with the piston seal and bronze cap missing from the actuator piston head.
The results of this testing indicates that this detail may have been overlooked in the original failure mode and effects analysis.
The aerodynamic effects of the thrust reverser plume on the wing, as demonstrated by simulation, has called basic certification assumptions in question.
Although no specific component malfunction was identified that caused uncommanded thrust reverse actuation on the accident airplane, the investigation resulted in an FAA determination that electrical and hydraulic systems may be affected.
As previously stated, the AD of August 15, required the deactivation of all electrically controlled B PW series powered thrust reversers until corrective actions were identified to prevent uncommanded in-flight thrust reverser deployment.
The condition of the left engine DCV which was recovered approximately 9 months after the accident, indicated that it was partially disassembled and reassembled by persons not associated with the accident.
Examination of the DCV indicated no anomalies that would have adversely affected the operation of the thrust reverser system.
The plug the investigation team found in the retract port of the DCV reference paragraph 1. However, the accident investigation team concluded that the plug a part used in the hydraulic pump installation on the engine was placed into the port after the accident by persons not associated with the investigation.
This determination was based on the fact that the plug was found finger tight which would indicate the potential for hydraulic fluid leakage with the hydraulic system operating pressure of psi applied.
Also, soil particles were found inside the valve body. However, their efforts were unsuccessful in that the procedure never led to identifying an anomaly.
When several attempts at the entire procedure were unsuccessful, Lauda personnel felt the need to continue troubleshooting efforts.
Boeing considers these removals and interchanges as not related to PIMU fault messages, ineffective in resolving the cause of the messages, and not per FIM direction.
Lauda maintenance records also indicate replacement and re-rigging of thrust reverser actuators. There was no further procedure or other guidance available in the Boeing FIM, and Lauda maintenance personnel made the decision to physically inspect the entire thrust reverser wiring harness on the engine and in the pylon.
If the message is cleared following a corrective action and does not reoccur on the next flight, when if it does reoccur, a new hour interval begins.
Therefore, Lauda was not remiss in continuing to dispatch the airplane and trouble shoot the problem between flights. No specific Lauda maintenance action was identified that caused uncommanded thrust reverser actuation on the accident airplane.
As a direct result of testing and engineering re-evaluation accomplished after this accident, Boeing proposed thrust reverser system design changes intended to preclude the reoccurrence of this accident.
In service B's were modified by incorporation of a Boeing service bulletin by teams of Boeing mechanics. The fleet modification was completed in February Design reviews and appropriate changes are in progress for other transport airplane.
The B design changes are based on the separation of the reverser deploy and stow functions by:. Adding a dedicated stow valve.
Adding new electric wiring from the electronics bay and flight deck to the engine strut. Critical wire isolation and protective shielding is now required.
Replacing existing reverser stow proximity targets with improved permeability material to reduce nuisance indications.
Adding a thrust reverser deploy pressure switch. The changes listed above for the B thrust reverser system address each of possible failure modes identified as a result of the investigation.
The design changes effectively should prevent in-flight deployment even from multiple failures. A diagram of the current at the time of the accident and new thrust reverse system is included in this report as appendix F.
Thrust reverser system reviews are continuing on other model series airplane. It was impossible to extract any information from the recorder.
Industry records indicate that investigative authorities have reported a similar loss of recorded data in several accidents that occurred both prior to and subsequent to the subject accident.
March 10, Dryden, Ont. There were some similar circumstances in each of the above mentioned accidents in that the crash site was located off airport property.
It was not possible for fire department vehicles to gain rapid access to the site. In each case, the FDR was involved in a ground fire which became well established and involved surrounding debris.
There does not appear to be a way to determine the exact duration of heat exposure and temperature level for the involved FDR in any of these accidents.
However, it has been recognized that ground fires including wood forest materials and debris continued in these instances for at least six to twelve hours.
The thermal damage to the tape recording medium was most probably the result of prolonged exposure to temperatures below the degree testing level but far in excess of the 30 minute test duration.
It is recommended that the airplane certification authorities and equipment manufacturers conduct research with the most modern materials and heat transfer protection methods to develop improved heat protection standards for flight data recorders.
Standards revisions should include realistic prolonged exposure time and temperature levels. The revised standards should apply to newly certificated FDR equipment and where practical through Airworthiness Directive action, to FDRs that are now in service.
The airplane was certificated, equipped and maintained, and operated according to regulations and approved procedures of the Republic of Austria.
The weather in the area was fair. There were no reported hazardous weather phenomena although lightning may have been present.
It is possible that the horizon was not distinguishable. The physical evidence at the crash site showed that the left engine thrust reverser was m the deployed position.
Examination of nonvolatile computer memory within the left EEC indicated that the engine was at climb power when the reverser deployed, engine thrust was reduced to idle with the reverser deployment, and the recorded Mach number increased from 0.
The actual maximum speed reached is unknown due to pressure measurement and recording uncertainties. The scatter of wreckage indicated that the airplane experienced in-flight breakup at a steep descent angle and low altitude.
Examination of the available wreckage revealed no evidence of damage from a hostile act, either from within the airplane or from the exterior.
Simulations of a 25 percent lift loss resulting from an in-flight deployment of the left engine thrust reverser indicated that recovery from the event was uncontrollable for an unexpecting flight crew.
From an airplane flight performance standpoint, questions remain unanswered regarding thrust reverser plume behavior at high Mach numbers and in-flight reverse induced airframe buffeting at high Mach numbers, and effects of inlet spillage caused by a reversed engine at high Mach numbers.
Thrust reverser system certification by the FAA required that the airplane be capable of continued safe flight and landing under any possible position of the thrust reverser FAR However, wind tunnel tests and data used in the simulation of this accident demonstrated that aerodynamic effects of the reverser plume in-flight during engine run down to idle resulted in a 25 percent lift loss across the wing.
Simulation of the event disclosed that the airplane was not capable of controlled flight unless full wheel and full rudder were applied within 4 to 6 seconds after the thrust reverser deployed.
However, no specific wire or component malfunction was physically identified that caused an uncommanded thrust reverser deployment on the accident airplane.
Testing identified hypothetical hydraulic system failures that could cause the thrust reverser to deploy. However, no specific component malfunction was identified that caused an uncommanded thrust reverser deployment on the accident airplane.
No specific Lauda Air maintenance action was identified that caused uncommanded thrust reverser deployment on the accident airplane.
The design changes recommended by Boeing and thereafter mandated by U. The specific cause of the thrust reverser deployment has not been positively identified.
The Aircraft Accident Investigation Committee also recommends that the United States Federal Aviation Administration revise the certification standards for current and future airplane flight recorders intended for use in accident investigation to protect and preserve the recorded information from the conditions of prolonged thermal exposure that can be expected in accidents which occur in locations that are inaccessible for fire fighting efforts.
Sound signatures identified as being produced by the engines were only visible when the power was advanced during the start of the takeoff roll.
No other definite engine signatures could be identified during any other portion of the recording.
Background "wind" noise in the cockpit can be heard to increase in intensity from thrust reverser deployment until the end of the recording.
This increase in background noise intensity is attributed to the aircraft's increasing airspeed during this span of time. The percentage of increase in the airspeed that the aircraft experienced during those final seconds of the recording could not be determined from the audio recording.
Also, during this time a noticeable modulation or vibration in the recorded sounds can be heard on the CVR recording. This anomaly in the recording was probably caused by the physical shaking of the recorder from airframe buffet.
Neither the United States National Transportation Safety Board nor the Boeing Company could demodulate this recorded vibration to obtain any meaningful data.
During the final seconds of the recording, several alarm or alert tones were heard on the CVR recording. National Transportation Safety Board along with the Boeing Company conducted a detailed investigation to document these tones.
There was insufficient information to form a definite conclusion as to the cause of these aural alerts. Pilot response to an upset condition.
Pilot response to an abnormal engine condition. Second actuation of the switch more than msec after first actuation.
The thrust reversers installed on the PW engines on the Boeing reverse only the fan airflow while the primary flow remains in the normal forward direction.
Thrust reversal is achieved by means of left and right hand translating fan sleeves containing blocker doors that block the fan flow redirecting it through stationary cascade vanes.
The translating sleeves are hydraulically actuated. Reverse thrust use is restricted to ground operation only, providing additional retarding force on the airplane during landings and refused takeoffs.
The FADEC results in the elimination of all engine control cables and the strut drum control box assembly.
Mechanical control features of the JT9D installation are replaced with electronic control. The Electronic Engine Control EEC uses throttle and reverser position inputs to allow commanded thrust levels forward or reverse.
The reverse thrust lever is lifted closing the Hydraulic Isolation Valve HIV switch which completes the circuit that opens the hydraulic isolation valve admitting hydraulic fluid to the thrust reverser system.
The isolation valve ports hydraulic fluid to the directional control valve DCV and also through the retract restrictor tee to the rod end of the actuators.
Further movement of the thrust lever closes the DCV switch thus allowing the DCV to port hydraulic fluid sequentially to the lock on the center actuator.
Hydraulic pressure build-up causes the lock piston to move and engages the lock lever pivot arm. Further motion of the piston separates the locking discs and fluid is ported directly to the head ends of the locking and non-locking actuators.
Linear movement of the actuator piston produces rotation of the high lead acme screw. The acme screw drives a gear train that is connected to the upper and lower actuators via flex drive shafts thus translating the reverser halves to the deploy position.
When both halves of the reverser reach the fully. To stow the reverser, the reverse thrust lever is returned to the fully down position thus opening the DCV switch which ports the actuator head end fluid to the return system.
Although the isolation valve switch on the thrust lever is also returned to the off stow position, auto restow switches operated by each reverser half of the reverser's translating sleeve remain closed and electrically hold the hydraulic isolation valve open until both halves are stowed.
The auto-restow circuit is automatically deenergized five 5 seconds thereafter. A two 2 second delay is used in this circuit to prevent nuisance illuminations.
Thrust Reverser Actuation System Description The thrust reverser is actuated by hydraulic power from three linear actuators attached to each translating sleeve.
The three actuators are synchronized by a flexible cable system contained within the hydraulic supply tubing. Supply and control of the hydraulic fluid to the actuators is by means of a hydraulic isolation valve, a directional control valve, and two flow restrictor orifice "T" connectors.
These three components are installed in the engine support strut. Hydraulic power is supplied to each reverser actuation system associated with the engine upon which the reverser is mounted.
When the solenoid is energized, the pilot valve is opened and fluid is ported to one end of an arming valve spool. This spool is spring biased to the closed position.
A pressure buildup of to psid is required to produce flow through the valve. A check valve is placed in the return port to prevent pressure surges from propagating back into the reverser's return system.
In addition to the de-energized and energized operating modes, the isolation valve has modes for inoperative dispatch and ground servicing.
For inoperative dispatch, a pin is inserted into the valve which prevents the valve arming spool from allowing fluid flow to the reverser actuators.
The DCV is dual-staged, with a solenoid operated pilot valve first stage and a hydraulic operated main valve second stage.
The DCV solenoid is powered through the DCV deploy switch which is mounted in a switch pack directly below the flight deck.
With the DCV solenoid deenergized stow mode and the HIV solenoid de-energized, the DCV main spool is spring and pressure biased to the stow mode and hydraulic pressure is applied to the rod end of the actuators only; the head end of the actuators are vented to return.
The actuators are maintained in the retracted stowed position. At 29 degrees of reverse thrust lever travel, the DCV switch is closed to deploy, thus energizing the DCV solenoid and allowing hydraulic fluid to pass through the first stage pilot valve.
Hydraulic pressure acting on a differential spool area then overcomes the spool spring force and shuttles the main valve spool to the deploy mode.
A damping orifice, located between the solenoid pilot valve and the main valve power spool, is used to reduce pressure spikes at the center actuator lock lever.
Flow Control System Orifice Tees The flow control system divides the incoming flow from the DCV to operate the two reverser sleeves on each engine as separate mechanisms operating simultaneously.
To accomplish this, the system incorporates flow restrictor tees in the extend and retract passages. During extension of the reverser, flow is routed through the extend restrictor tee to the actuator head ends.
Equal pressure is developed in both head and rod end cavities of the actuators. Reverser extension is achieved by having a two-to-one actuator piston area differential favoring extension.
The returning flow from the actuator rod ends is routed through the retract restrictor tee and ports to the PRESS B port of the directional control valve.
Actuators The six actuators used to operate each engine's thrust reverser sleeves are hydraulically powered. Actuator movement in the extend direction is produced by connecting both head and rod end cavities to the source of flow thus providing an extension force equal to the supply pressure acting over the difference between head and rod end areas.
Actuator movement in the retract direction is produced by connecting the rod end cavity to supply and the head end cavity to return. The linear movement of the actuator piston produces rotation of an acme screw that is installed concentric within the piston rod.
The piston rod is prevented from rotational motion relative to the actuator body by the gimbal mount of the actuator and pinned attachment of the rod end.
Rotation of the acme screw drives the synchronization gear train. The synchronization gear trains of adjacent actuators are connected by flexible cables that are encased within the hydraulic tubing that connects the head end cavities of these actuators.
A square end drive on each end of the flexible cables inserts into the worm gear of the synchronization gear train to complete the mechanical connection.
As the actuators extend, fluid flow to the head ends is provided by one-half of the volume coming from the fluid source and one-half the volume coming from the restrictor tee of the flow control system and returned to port PRESS B of the DCV.
Fluid flow to and from the rod end cavity is ported through the snubbing ring. When the actuator is extending, outflow passes to the hydraulic fluid fitting on the actuator rod end.
Snubbing begins when the snubbing skirt on the piston rod enters the gap between the piston rod and the snubbing ring. The reverser retract cycle is not snubbed because the retracting velocities are lower and there is no driving aerodynamic loads.
Locking Actuators Each half sleeve for each engine reverser is translated with three hydraulic linear actuators. The center actuator on each half sleeve incorporates a locking mechanism that functions by engagement of two serrated discs.
This engagement directly prevents rotation of the synchronizing gear train that mechanically interconnects the three actuators. One disc is keyed to the acme screw in the actuator and rotates when the actuator is translating.
The other disc is non-rotating, splined to the actuator barrel, and is actuated linearly along the spline by a helical.
As the center actuator nears the stowed position during retraction the helical lock spring becomes compressed forcing the splined, non-rotating disc against the rotating disc.
This causes the two discs to ratchet until the actuator piston bottoms. The center actuator is locked against extension by serration engagement which prevents acme screw rotation and hence piston movement.
During retraction, the return flow from the actuator bead end bypasses the lock piston through a check valve and the preload spring holds the lock piston in the locked position.
The spring bias of the preload spring also prevents pressure surges from inadvertently unlocking the serrated disks while the reverser is stowed.
Thrust Reverser Position Feedback System The thrust reverser feedback system provides the EEC with an indication of the thrust reverser sleeve positions as measured at the center locking actuators.
There are two separate electrical inputs, outputs, moveable armatures, etc. The two movable armatures are joined together and are driven by a single mechanical input.
As the actuators are extended or retracted, the armatures are inserted into or withdrawn from the LVDT stator, respectively.
This is included in the system in the event of a mechanical failure of the feedback linkage from the center locking hydraulic actuators.
Six switches must all be closed to obtain hydraulic flow in the reverser system for normal reverser system for normal reverser operation.
Three switches must be closed to complete the circuit to the isolation valve. Either one of two auto-restow sensors, independent of the preceding six switches, initiate or maintain reverser operation any time either reverser half is not stowed.
Reverser operation is initiated by energizing the solenoid that opens the isolation valve. Fire Switches Operating the fire switches will remove electrical power from the isolation valve and the directional control valve solenoids.
Isolation Valve Switch The isolation valve switch is a micro switch mounted near the hinge point of the thrust reverse lever. The switch is activated by a contoured surface at the hinge of the lever.
The switch closes at any time the thrust reverse lever is lifted more than 10 degrees. The switch is activated by a contoured surface on the switch cam via a follower and roller assembly.
The switch closes and energizes the DCV any time the thrust reverse lever is lifted more than 29 degrees. Auto Restow Sensors Two proximity sensors, one for each reverser half, are located on the nacelle torque box structure at the forward end of the reverser cascade near the reverser's center actuator.
The target elements for the switch sensors are located on the translating sleeve. The sensors are adjusted to close when the reverser sleeve moves from the fully stowed position.
The stow relay is energized to complete an electrical circuit to the isolation valve. Since the reverser hydraulic power must remain available until the reverser is fully stowed during the stow cycle, a 5 second time delay following the sensed reverser stowed position is incorporated in the Proximity Switch Electronic Unit PSEU logic for the restow circuit.
System Separation The electrical circuit controlling the left engine thrust reverser is separated from the right engine.
Separate power sources, circuit breakers, switches, wires, and relays through to separate isolation valves are used.
The individual reverser wire bundles are routed separately from each other. The auto-restow proximity sensors are connected to separate sections of the proximity switch electronic unit PSEU.
The control circuits to the HIV and DCV solenoids are electrically separated from the indication circuit on each engine. Proximity Sensor The auto-restow proximity sensors are excited by an electronic circuit in the PSEU mounted in the electrical rack.
The circuit and power source for the left thrust reverser restow sensors are separate from that of the right engine reverser.
Reverser unlock is indicated by "REV" in amber color. In full deploy "REV" changes to green. A two-second time delay is used with this isolation valve indication to remove nuisance warnings.
Reverser Unlocked Indication The reverser unlocked indication is activated by either of two proximity switches located one on each lock housing of the center actuators.
The "REV" amber indication occurs anytime either lock is unlocked. The proximity switch is activated by movement of a target arm attached to the lock actuator's pivot shaft.
Full Reverse Indication The full reverse indication is controlled by two proximity switches which are connected so that the "REV" green indication occurs only when both reverser halves reach the fully deployed position.
In the event that amber and green are signalled simultaneously, the amber signal prevails. L R REV ISLN VAL caution indicates that a malfunction exists that may result in a reverser deployment if the thrust reverse lever is lifted in flight, or that on reverser may not deploy when commanded on the ground.
The indication is required because the pilot may not be able to detect the interlock failure to block thrust lever motion during normal thrust reverser deployment.
A status message will be sent to EICAS alerting the crew of the lack of interlock for the landing aid the next dispatch. System Separation The electronic circuits operating the proximity switches and reverser indication are located in the proximity switch electronic unit module PSEU mounted in the electronic rack.
Complete separation is maintained between the left and right engine circuits with separate power sources, circuit breakers, wire, and relays.
Power which is generated by the HMG is transferred to the right and left thrust reversers via the standby and battery busses.
If normal power is recovered during flight such that both main busses are energized, the HMG shuts down to allow normal system operation.
The main function of the EEC is the scheduling of fuel flow, stator vanes and bleed valves to control the thrust and performance of the engine as a function of the thrust lever position.
The EEC is configured as a dual channel system with independent inputs to and outputs from each channel. The reverser position is provided as an electrical signal to each EEC channel by two independent position sensing circuits containing linear variable differential transducers LVDT.
The LVDT's sense each sleeve position from the center actuators. Each channel's output of one dual LVDT is connected in series electrically to the corresponding channel's output of the dual LVDT mounted on the other sleeve's locking actuator.
The LVDT electrical inputs for each channel are wired in parallel. These series connected LVDT outputs provide an indication of the average reverser sleeve position to each channel primary and secondary of the EEC, while maintaining electrical separation of the EEC channels.
Each EEC channel provides a discrete output which energizes the interlock actuator relay. Thrust Limiting Function This function compares the thrust commanded by the pilot TRA to the position of the thrust reverser sleeves.
The limiting function is incorporated to ensure that thrust is in the direction of the command. This function is invoked under two circumstances, the first occurs when the direction of commanded thrust has just changed and the reverser is in transit to the commanded position.
Mechanical interlocks are incorporated to prevent the pilot from commanding reverse thrust above idle until the thrust reverser is at a prescribed position.
Thrust limiting in the EEC, during normal operation, provides a second level of protection against high thrust in the uncommanded direction.
Thrust limiting will also be invoked in the case of an inadvertent departure of the thrust reverser from the commanded position.
The EECs thrust limiting function provides an independent system to reduce the engines thrust until the sleeve position agrees with the TRA command.
July 3, In reply refer to: A through Honorable James B. All passengers and 10 crewmembers on board were fatally injured in the accident.
The positions of the left engine thrust reverser actuators along with data from the electronic engine control EEC and the cockpit voice recorder CVR indicate that the left engine thrust reverse system deployed while the airplane was at approximately.
The preliminary evidence suggests that the reverse event was recognized by the flightcrew but that the airplane departed controlled flight, accelerated past the maximum operating velocity, and experienced an in-flight structural breakup.
Indications of an in-flight fire prior to the breakup have not been found. However, during the breakup, a large explosion was witnessed and burning debris fell to the ground.
The accident airplane was equipped with Pratt and Whitney PW series engines. The Boeing Airplane Company provides an electro-hydraulic thrust reverse system in these airplanes to redirect engine fan bypass airflow to aid in stopping the airplane on the ground.
The thrust reverse system contains logic switching devices that are designed to prevent in-flight deployment caused by a component failure or flightcrew action.
These engines also incorporate EEC devices. One function of the EEC is to reduce engine rpm to idle in the event of an inadvertent reverser deployment.
Although a reduction in reverse thrust is beneficial, it does not occur immediately because of the time delay while the engine spools down.
The thrust reverse system of the PW series engines installed in Boeing airplanes incorporates a hydraulic isolation valve HIV and a directional control valve DCV in the engine pylon.
The CVR revealed that the flightcrew observed the REV ISLN caution light illuminated about 9 minutes prior to the reverser deployment on the accident airplane and a crewmember observed that the light came on repeatedly.
The flightcrew discussed the Boeing Quick Reference Flight Handbook QRH information which states that if this caution light is illuminated, additional systems failures may cause inflight deployment.
The thrust reverse system is designed so that the HIV provides a safeguard against deployment caused by a DCV failure. The system is designed so that the HIV will open to provide pressure to the reverser system in flight to restow the thrust reverser if it is not fully closed.
The valve can also open when certain faults exist in the system logic. The HIV normally opens when the airplane lands and the reverse system is used.
A DCV failure might then be apparent when the translating cowl does not stow properly. While information provided by the manufacturer indicates that other Boeing airplanes have experienced 'REV ISLN' caution light illuminations during flight, there have been no prior indications of DCV failure or uncommanded thrust reverser extensions.
The hydraulic thrust reverse actuators from the left engine of the accident airplane were found in the deployed position and no pre-existing faults were evident.
Hydraulic power for the actuators can come only through the DCV located in the pylon, which is a high vibration environment.
The left engine DCV has not been found and thus could not be examined for malfunction. It was located in the pylon near the point where the pylon separated, from the airplane.
However, a failure mode and effects analysis for the thrust reverser system has revealed failure modes in the DCV that could cause an uncommanded reverser deployment after an opening of the HIV.
The Safety Board has been provided with data from Boeing indicating that flight control has been demonstrated on the Boeing with one engine in the reverse idle position at knots IAS; however, the Board has been informed that such testing has not been performed at higher speeds or at higher engine thrust levels.
The Safety Board is concerned about the potential severity of airframe buffeting, aerodynamic lift loss, and subsequent yawing and rolling forces which may occur at the airspeed and engine thrust levels that existed when the reverser deployed in the accident flight.
The Safety Board is also concerned that Boeing flightcrew emergency procedures may not provide appropriate and timely guidance to avoid loss of flight path control in the event that the reversers deploy in flight.
Pending completion of actions taken to assure acceptable reliability of the thrust reverse system, the Safety Board believes that flight crew procedures in response to a 'REV ISLN" light while airborne should include actions to attain appropriate combinations of altitude, airspeed, and thrust settings which will minimize control difficulties in the event of subsequent reverser deployment.
Furthermore, consideration should be given to the development of emergency procedures which would include pulling the fire handle in the event that the reverser does deploy.
This would immediately remove fuel, and hydraulic and electrical power to the affected engine. The Safety Board also believes that flightcrews should be forewarned that an in-flight deployment of a thrust reverser may result in significant airplane buffeting, yawing, and rolling forces.
Conduct a certification review of the PW engine equipped Boeing airplane thrust reverser systems to evaluate electrical and mechanical anomalies and failure modes that can allow directional control valve pressure to be applied to the reverser EXTEND port.
The certification review should include subjecting the valve to the engine's vibration spectrum concurrent with introduction of intermittent pressure spikes to the valve pressure port.