Unfall lauda

unfall lauda

1. Aug. Nach seinem Unfall: Niki Lauda im Jahr bild: imago Niki Lauda dominiert als Weltmeister die FormelSaison. Doch beim GP auf dem. 8. Aug. Walter Klepetko, Leiter der Thoraxchirurgie, der Lauda mit seinem durch einen FormelUnfall im Jahr vorgeschädigten Lunge auf. Aug. FormelStar, Funktionär, Airline-Besitzer. Aber vor allem: Kämpfer!Niki Lauda ( 69) ist ohne jeden Zweifel eine der größten.

Unfall Lauda Video

Formel1 1976 GPvD Nürburgring (hyperborean.nu)

The audible fire warning system in the cockpit was silent. The absence of soot on the cabin outflow valve and in the cargo compartment smoke detectors indicates that no in-flight fire existed during pressurized flight.

Evidence indicates that the fire that developed after the breakup resulted from the liberation of the airplane fuel tanks.

No shrapnel or explosive residue was detected in any portion of the wreckage that was located. Evidence of an explosion or fire in the sky was substantiated by witness reports and analysis of portions of the airplane wreckage.

Although it is possible in some cases that some "in-air" fire damage was masked by ground fire damage, only certain portions of the airplane were identified as being damaged by fire in the air.

These include the outboard wing sections and an area of right, upper fuselage above the wing. Evidence on the fuselage piece of an "in-air" fire include soot patterns oriented with the airstream and the fact that the piece was found in an area of no post-crash ground fire.

Evidence of an "in-air" fire on the separated outboard portions of the right and left wings include that they were found in areas of no ground fire, yet were substantially burned.

The separated right wing portion had been damaged by fire sufficiently to burn through several fuel access panels. In addition, one of the sooted fractures on the right wing section was abutted by a "shiny" fracture surface.

These fracture characteristics show that the separation of the right wing section had preceded its exposure to fire or soot in the air, followed by the ground impact that produced the final, "shiny" portion of the fracture.

Generally, it appears that fire damage was limited to the wings and portions of the fuselage aft of the wing front spar except for the left mid-cabin passenger door.

Likewise, many areas of the fuselage aft of the wing front spar were devoid of fire damage. This is further indication that the airplane was not on fire while intact, but started burning after the breakup began.

The absence of any fire damage on the empennage indicates that it had separated prior to any in-air fire. The sooting documented on the left mid-cabin passenger door is unique in that the fuselage and frame around the door were undamaged by fire or soot.

Even the seal around the door appeared to be only lightly sooted. The door was found in an area of no ground fire, indicating that the door was sooted before ground impact.

The sooting on the door, but not on the surrounding structure, may have resulted as the door separated from the fuselage during the breakup and travelled through a "fire ball" of burning debris.

It is not known why the door seal did not exhibit the same degree of sooting as the door itself, although it is possible that the soot would not adhere to the seal as well as to the door.

These efforts yielded erroneous results because the simulators were never intended for such use and did not contain the necessary performance parameters to duplicate the conditions of the accident flight.

NTSB requested the Boeing Commercial Airplane Group to develop an engineering simulation of in-flight reverse thrust for the conditions thought to have existed when the left engine thrust reverser deployed in the accident flight.

As previously stated, the flight data recorder FDR tape in the accident airplane was heat damaged, melted, and unreadable due to post-crash fire.

Flight conditions were therefore derived from the best available source, post-accident readout of the left engine EEC non-volatile memory parameters.

Test conditions were proposed by Boeing and accepted by the participants as follows: The left engine thrust reverser was configured to provide reverse thrust effect at the start of reverse cowl movement rather than phased to cowl position.

The right engine was set up to be controlled by the pilot through the throttle handle. Tests were run with pilot commanded right engine throttle cutback to idle following the reverser deployment on the left engine.

Tests were repeated with no throttle cutback on the right engine. The autopilot was engaged in single channel mode for all conditions.

Upon initiation of pilot recovery action, the autopilot. The autopilot does not operate the rudder under the conditions experienced by the accident airplane.

The autopilot operates the rudder only while in the "autoland" mode of flight. However, it was not considered to be significant. The left engine electronic control indicates that the thrust reverser deployed in the accident flight at approximately 0.

There were no high-speed wind tunnel or high-speed flight test data available on the effect of reverse thrust at such an airspeed.

To be suitable for use in the engineering simulation, in-flight reverse thrust data were needed for an airplane of similar configuration to the B This similarity was essential because the intensity and position of the reverse thrust airflow directly affects the controllability of the airplane.

Airplanes with wing-mounted engines such as the DC-8, DC, B and B have experienced in-flight reverse thrust, and according to Douglas Airplane Company, all models of the DC-8 including those airplanes retrofitted with high-bypass fan engines were certificated for the use of reverse thrust on the inboard engines in flight.

Although the B has wing-mounted engines, it also has longer engine pylons which place the engines farther ahead and below the leading edge of the wing compared to the B Available in-service data suggests that the farther the engine is located from the wing, the less likely its reverse thrust plume will cause a significant airflow disruption around the wing.

The B has wing mounted engines, however, its reverser system is located in the rear of the engine, below and behind the wing leading edge, also making it less likely to affect wing lift.

In the case of in-flight reverse thrust on large three or four engine airplanes, each engine produces a smaller percentage of.

Based on engineering judgement the lower proportion of thrust and resultant airflow affects a smaller percentage of the wing, and therefore the effect of reverse thrust is less significant on a three or four engine airplane than on a two engine airplane.

The mechanical design and type of engine is also important in the event of in-flight reverse thrust.

The B's engines are high-bypass ratio turbofans, with reverser systems which employ blocker doors and cascades to redirect airflow from the N 1 compressor fan blades.

On large twin-engine transport airplane, the thrust reverser cascades are slightly below and in front of the wing. At high thrust levels, the plume of thrust from the reverser produces a yawing moment and significantly disrupts airflow over the wing resulting in a loss of lift over the affected wing.

The loss of lift produces a rolling moment which must be promptly offset by coordinated flight control inputs to maintain level flight.

The yaw is corrected by rudder inputs. 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. Please activate JavaScript within your browser settings.

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Austrian Date of Birth: Parmalat Racing Team - 14th Formula One: Parmalat Racing Team - 4th Formula One: Scuderia Ferrari - Champion Formula One: Scuderia Ferrari - 2nd Formula One: Scuderia Ferrari - 4th Formula One: Sorry, es scheint so, als sei die Version deines Browser zu alt!

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He has arranged for sponsors to use the cap for advertising. With Lauda out of the contest, Carlos Reutemann was taken on as his replacement.

Lauda missed only two races, appearing at the Monza press conference six weeks after the accident with his fresh burns still bandaged.

He finished fourth in the Italian GP , despite being, by his own admission, absolutely petrified. F1 journalist Nigel Roebuck recalls seeing Lauda in the pits, peeling the blood-soaked bandages off his scarred scalp.

He also had to wear a specially adapted crash helmet so as to not be in too much discomfort. In Lauda's absence, Hunt had mounted a late charge to reduce Lauda's lead in the World Championship standings.

Hunt and Lauda were friends away from the circuit, and their personal on-track rivalry, while intense, was cleanly contested and fair.

Lauda qualified third, one place behind Hunt, but on race day there was torrential rain and Lauda retired after two laps.

He later said that he felt it was unsafe to continue under these conditions, especially since his eyes were watering excessively because of his fire-damaged tear ducts and inability to blink.

Hunt led much of the race before his tires blistered and a pit stop dropped him down the order. He recovered to third, thus winning the title by a single point.

Lauda's previously good relationship with Ferrari was severely affected by his decision to withdraw from the Japanese Grand Prix, and he endured a difficult season , despite easily winning the championship through consistency rather than outright pace.

Lauda disliked his new teammate, Reutemann, who had served as his replacement driver. Lauda was not comfortable with this move and felt he had been let down by Ferrari.

It suffered from a variety of troubles that forced Lauda to retire the car 9 out of 14 races. Lauda's best results, apart from the wins in Sweden and Italy after the penalization of Mario Andretti and Gilles Villeneuve, were 2nd in Montreal and Great Britain, and a 3rd in the Netherlands.

As the Alfa flat engine was too wide for effective wing cars designs, Alfa provided a V12 for It was the fourth 12cyl engine design that propelled the Austrian in F1 since Lauda's F1 season was again marred by retirements and poor pace, even though he won the non-championship Dino Ferrari Grand Prix with the Brabham-Alfa.

After that, Brabham returned to the familiar Cosworth V8. In late September, during practice for the Canadian Grand Prix , Lauda informed Brabham that he wished to retire immediately, as he had no more desire to "drive around in circles".

Lauda, who in the meantime had founded Lauda Air, a charter airline, returned to Austria to run the company full-time. In Lauda returned to racing.

After a successful test with McLaren , the only problem was in convincing then team sponsor Marlboro that he was still capable of winning.

Lauda proved he was when, in his third race back, he won the Long Beach Grand Prix. Before the opening race of the season at Kyalami race track in South Africa , Lauda was the organiser of the so-called "drivers' strike"; Lauda had seen that the new Super Licence required the drivers to commit themselves to their present teams and realised that this could hinder a driver's negotiating position.

The drivers, with the exception of Teo Fabi , barricaded themselves in a banqueting suite at Sunnyside Park Hotel until they had won the day.

Some political maneuvering by Lauda forced a furious chief designer John Barnard to design an interim car earlier than expected to get the TAG-Porsche engine some much needed race testing; Lauda nearly won the last race of the season in South Africa.

Lauda won a third world championship in by half a point over teammate Alain Prost , due only to half points being awarded for the shortened Monaco Grand Prix.

Initially, Lauda did not want Prost to become his teammate, as he presented a much faster rival. However, during the two seasons together, they had a good relationship and Lauda later said that beating the talented Frenchman was a big motivator for him.

Lauda won five races, while Prost won seven. However, Lauda, who set a record for the most pole positions in a season during the season, rarely matched his teammate in qualifying.

Despite this, Lauda's championship win came in Portugal , when he had to start in eleventh place on the grid, while Prost qualified on the front row.

Prost did everything he could, starting from second and winning his 7th race of the season, but Lauda's calculating drive which included setting the fastest race lap , passing car after car, saw him finish second behind his teammate which gave him enough points to win his third title.

His second place was a lucky one though as Nigel Mansell was in second for much of the race. However, as it was his last race with Lotus before joining Williams in , Lotus boss Peter Warr refused to give Mansell the brakes he wanted for his car and the Englishman retired with brake failure on lap As Lauda had passed the Toleman of F1 rookie Ayrton Senna for third place only a few laps earlier, Mansell's retirement elevated him to second behind Prost.

After announcing his impending retirement at the Austrian Grand Prix , he retired for good at the end of that season. After qualifying 16th, a steady drive saw him leading by lap However, the McLaren's ceramic brakes suffered on the street circuit and he crashed out of the lead at the end of the long Brabham Straight on lap 57 when his brakes finally failed.

He was one of only two drivers in the race who had driven in the non-championship Australian Grand Prix , the other being World Champion Keke Rosberg , who won in Adelaide in and would take Lauda's place at McLaren in ,.

In Lauda returned to Formula One in a managerial position when Luca di Montezemolo offered him a consulting role at Ferrari. Halfway through the season Lauda assumed the role of team principal of the Jaguar Formula One team.

The team, however, failed to improve and Lauda was made redundant, together with 70 other key figures, at the end of Set in the golden era of Grand Prix Racing '1' tells the story of a generation of charismatic drivers who raced on the edge, risking their lives during Formula 1's deadliest period, and the men who stood up and changed the sport forever.

In Niki Lauda survived one of the most famous crashes in Formula One history. Using previously unseen footage, Lauda: The Untold Story explains what happened on that fateful, and near fatal day at the Nurburgring, then follows Lauda's courageous journey to recovery culminating in a miraculous comeback in Monza just weeks later.

The film also investigates the impact that his crash had not just on his own life but on the sport as a whole, looking at the safety developments from the s to the present day.

The Untold Story is a must-see for all motor-sport fans. I don't know why this got a cinema release, the quality of editing and production standards are barely up to what you can watch on TV.

The documentary contains quite interesting interviews and some new footage but the final half of it resembles an instruction video about F1 safety technology that ads nothing to the story of Lauda.

The editing and chapter breaks make it hard to follow the timeline of events. The film also delves into too much of the history of F1 racing. In all the director spends far too little time telling the story of Lauda and frequently goes of on tangents that have little to do with Lauda's story.

The film is also regularly inter-sped with an unnecessary narrated voice-over and loud music. The director should watch Senna for a masterclass in documentary making as his film pales in comparison.

Start your free trial. 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.

The certification review should also determine the adequacy of the thrust reverser system safeguards when the hydraulic isolation valve is open to prevent uncommanded thrust reverser extensions.

Class I, Urgent Action A Amend the Boeing Flight Operations Manual on aircraft powered by the PW series engine to include in the section, "Reverser Isolation Caution Light," a warning that in-flight reverser deployment may result in severe airframe buffeting, yawing, and rolling forces.

Class I, Urgent Action A Pending completion of a certification review of the thrust reverser system, establish operational procedures to be followed upon illumination of the Reverse Isolation Caution Light REV ISLN that will enhance the controllability of the PW powered Boeing should a secondary failure result in the in-flight deployment of a thrust reverser.

Actions should be taken to achieve an appropriate. Also consider the inclusion of a procedure to pull the fire handle if this occurs.

In lieu of implementation of revised operational procedures, operators may be directed to deactivate thrust reversers until the certification review is completed and the reliability of the system can be adequately assured.

Class I, Urgent Action A Examine the certification basis of other model airplanes equipped with electrically or electro hydraulically actuated thrust reverse systems for appropriate safeguards to prevent inflight deployment of reversers and ensure that operating procedures are provided to enhance aircraft control in the event an of inadvertent in-flight reverser deployment.

To airworthiness authorities of countries having operators of Boeing Model , , , and airplanes. This letter represents a continuation of the series of letters describing the FAA's actions in response to a recent accident which apparently resulted from an uncommanded Inflight thrust reverser deployment on a Boeing Model ER airplane.

The FAA is cooperating in this investigation and, in addition, is reviewing thrust reverser certification philosophy and the design of current thrust reversers.

There will be future actions taken by the FAA to assure the safety of thrust reverser systems. The rules for thrust reverser certification assume that inflight reverser deployments will occur and they require that such deployments not result in an unsafe condition.

Traditionally, this has been demonstrated by tests conducted at relatively low speed and thrust conditions supported by analytical extrapolations to all flight conditions.

Service experience on many airplane models has included inflight deployments which were controllable and appeared to validate these certification procedures.

These procedures were applied to the certification effort, and indicated that an inflight reversal was a controllable event. The recent accident calls these certification assumptions into question.

It is possible that modern high bypass engines combined with more efficient thrust reversers have resulted in aircraft which require a new thrust reverser certification philosophy.

Inflight reversal, under certain flight conditions, may now be an event similar in magnitude to certain primary flight control failures which must be prevented to avoid loss of the aircraft.

The Boeing Company is in agreement with the need to upgrade the level of safety of thrust reverser systems, and has been cooperating with the FAA in a review of all of their thrust reverser installations.

This includes system design philosophy and system design details. This review, of course, began with the due to the recent accident.

Review of the thrust reverser installations in other Boeing airplanes has been proceeding and is now to a point where some future actions can be defined.

These actions include interim actions to assure the safety of thrust reversers and long term design changes and retrofit to bring thrust reverser systems up to safety level of primary flight controls.

This review, will discuss each Boeing airplane modal separately, and will present plans for both interim and final action.

These are as follows: At present, all thrust reverser systems on these air planes are deactivated due to the issuance of airworthiness directive AD T, dated August 23, Boeing is at present studying several proposals for interim design changes, which would assure an increased level of safety for this thrust reverser system, thus permitting reactivation of these thrust reversers pending a final revision to the design.

Boeing intends to present their interim design change proposal to the FAA during the week of September 9, , and it is anticipated that service bulletins would be available for FAA review and approval during the week of September 23, If it is determined by the FAA that the proposal provides adequate safeguards, it is the intention of the FAA to mandate this design change by AD action, and permit reactivation of the affected thrust reverser systems.

When a final design change has been approved, it in turn will be mandated by ad action, it is anticipated that these design changes will reduce or eliminate the requirement for repetitive tests and inspections of the thrust reverser system.

At present, operation of these airplanes with active thrust reverser systems is permitted. It is anticipated that certain repetitive system tests and inspections will be mandated by AD action.

The service bulletins necessary for these tests and inspections have already bean approved by the FAA. In addition, the electrical wiring for these airplanes is being examined for adequacy with respect to system separation and hot short protection.

At the completion of this investigation, it is expected that a final design change will be generated, which will reduce or eliminate the requirement for repetitive tests and inspections of the thrust reverser system.

Since these thrust reverser systems employ mechanically actuated directional control valves, it is felt that they do not possess the same potential for inflight reversal as those systems listed above.

This assumption is further supported by a trouble free service history to date with respect to uncommanded inflight deployments.

A comprehensive investigation of the hydraulic system is in progress, and any AD action will depend upon the results of this investigation.

When a final design change is approved, It will be mandated by AD action. As an interim action, the FAA is issuing an immediate adopted AD the week of September 9, , which mandates initial and follow-on thrust reverser electrical system checks and replacement of those DCV solenoid valves which are susceptible to the contamination failure.

A copy of the AD is included with this letter. The Boeing Company is currently evaluating long term thrust reverser system configuration changes which could be terminating action to all or part of the repetitive electrical system inspections.

Boeing Model series airplanes powered by Rolls Royce RB engines employ a different hydraulically actuated thrust reverser design.

This system is not susceptible to the contamination failure cited in the AD. Design changes are being developed by Boeing to improve the reverser system.

Boeing Model engine thrust reverser systems: Any applicable system improvements identified for the systems will be required on the in the long term.

No immediate actions are being taken on the because aerodynamic differences between the and the twin-engine airplanes result in adequate controllability with a reverser deployed.

Nevertheless, the FAA believes and Boeing agrees that inflight thrust reversals are undesirable, and all design improvements identified for the thrust reverser system will also be incorporated on airplanes.

Boeing has indicated that it plans to release system check service bulletins for the thrust reverser systems in the near future.

The FAA recommends that any Boeing-provided system cheeks be performed, but there are no current plans to release airworthiness directives requiring the performance of the system checks contained in these service bulletins.

While there are no plans for FAA action as of this date, results of these investigations may require that steps be taken to incorporate features or activities consistent with actions taken on other models.

In closing, we would like to point out that, in addition to the above, you should be aware that the Transport Airplane Directorate is conducting a Design review of the thrust reverser installations on other large jet transports manufactured by McDonnell-Douglas, Airbus Industries, Lockheed, etc.

As a result of that review, design changes may be required in the future. We request that you ensure that this letter is made available to airline flight departments and to all pilots of the above Boeing airplanes, to keep them fully apprised of the progress of this investigation.

Sincerely, original signed by Leroy A. Comments of the accredited representative of the United States of America were brief; and incorporated in the Final Report.

Comments of the accredited representative of Austria are appended. Brief edit items were incorporated in the Final Report. Comments on airline maintenance activities and the calls for further testing and analysis of the effects of reverser deployment and reexamination of the Dispatch Deviation Guide are provided to enlighten the reader.

These items were not included in the Final Report or Recommendations. The Accredited Representative of the Republic of Austria, whilst agreeing that this report is a fair record of the investigation, regrets that the report was unable to form any conclusion as to the reason for the uncommanded thrust reverser deployment which was the fundamental cause of the accident.

Whilst acknowledging the modifications package developed for aircraft similar to the Lauda machine and the recommendation for design reviews of all other aircraft certificated for ground-use only reverser systems, the lack of knowledge about the aerodynamic effects of deployment at high Mach numbers and Indicated Air Speeds should not be allowed to persist.

Accordingly it is felt that the report should call for further analysis and testing to be accomplished on the effects of reverser deployment throughout the flight envelope on aircraft of similar configuration to the Boeing In addition, it is noted that the requirements of FAR I am concerned by the apparent lack of analysis of the Cockpit Voice Recorder, being the only continuous record of the accident event in the.

There appears to be no attempt to interpret anything other than the cockpit speech although the recording contained considerably more recorded intelligence which, if anlalysed in-depth, may have yielded information about the crew's and the aircraft's behaviour following the inadvertent deployment.

I am also of the opinion that the Boeing Company's interpretation of their own Dispatch Deviation Guide requirements should be reexamined.

A repetitive EEC fault message that continues for some hours despite rectification actions is clearly not responding to these actions and yet could theoretically continue indefinitely as long as it does not manifest itself during the hour period allowed by the Dispatch Deviation Guide.

The following changes are proposed to be incorporated in the Final Report as they are written bold Italic , other comments should cause a more detailed explanation in the report:.

Page 2 Line Page 3 Line 7: The pilot-in-command, male, age 48, December 19, , valid until December 31, The co-pilot, First officer, male, 41 years of age, Civil Aviation of Austria issued April 24, Valid until October 24, Page 4 Line Post accident interrogation of the EEC non volatile memory, which dated to April 27 indicated a significantly higher number of similar messages occurred than were recorded in the documentations.

There was no radar recording of the accident flight available. Page 7 Line We feel the need to notify wind in velocity and direction in ft altitude steps under this headline or in the wreckage diagramme.

We feel the need to provide more evidence on the nature and the extent of the inflight fire or how the conclusion came up, that it occurred after the inflight break up.

Page 14 Line Page 15 Line 1: Page 16 Line 3:

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