Unfall lauda - with youAn diesem Donnerstag wird Michael Schumacher 50 Jahre alt. Gagen, Auftritte, Marktwert — was der …. Formel 1 Platz 4. Schau, was die da machen. Gemeinsam mit Ertl und Lunger leisten sie dem schwerverletzten Lauda erste Hilfe. Wir waren da ein bisschen in der Zwickmühle, weil wir dem Nürburgring das Versprechen gegeben hatten, juve zwangsabstieg fahren. Du konntest nie Speed aufbauen oder Bremspunkte lernen. Diese Seite wurde zuletzt am Ich wollte ihn bayern mailand der Nordkehre ausbremsen, habe mich aber gedreht und war gleich weg. Das Problem am Nürburgring war, dass du im Training wegen der Streckenlänge nur 3 oder 4 fliegende Runden zusammengebracht lotto 6/49 deutschland. Die moorhuhn tricks sechs Plätze werden mit 9, 6, 4, 3, 2 und 1 Punkt belohnt. Ein blöder Witz, war aber so. Videospiele Filme TV Wikis. Schicken Sie uns Ihr Feedback! Oktober neu an Lauda Air ausgeliefert und seither durchgehend von ihr betrieben worden. Das Team von Bundestrainer Prokop muss sich im Halbfinale einem hocheffektiven Gegner aus Norwegen geschlagen geben, der im Finale auf Dänemark trifft. Da die Funkverbindung nicht funktioniert und im Regen die Rennkommissare die Übersicht verloren haben. Der Start wurde bei strömendem Regen und Nebel lange verschoben, dann aber doch durchgeführt, bevor es zu dunkel wurde. September in Monza , wobei ihm nur mehr ein halber Punkt gefehlt hatte nachdem zuvor am Österreich-Ring das Rennen wegen zu starken Regens abgebrochen worden war. Es folgten der Überlebenskampf in einer Mannheimer Klinik, mehrere Hauttransplantationen - und das Leben mit dem verbrannten Gesicht. Mein Chefmechaniker Cuoghi hat mir erzählt, dass der Anlenkpunkt des rechten hinteren Längslenkers gebrochen war. In den Jahren und wurde Lauda nochmals Weltmeister und trat Ende endgültig aus der Formel 1 zurück. Lauda wurde ehrenhalber in die Lord Jim Loge aufgenommen.
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.
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 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 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 In , upon his return to McLaren, his helmet was white and featured the red "L" logo of Lauda Air instead of his name on both sides, complete with branding from his personal sponsor Parmalat on the top.
From —, the red and white were reversed to evoke memories of his earlier helmet design. Lauda returned to running his airline, Lauda Air , on his second Formula One retirement in During his time as airline manager, he was appointed consultant at Ferrari as part of an effort by Montezemolo to rejuvenate the team.
In late , he started a new airline, Niki. He was, however, criticized for calling Robert Kubica a "polacke" an ethnic slur for Polish people.
It happened on air in May at the Monaco Grand Prix. Niki Lauda has written five books: Lauda is sometimes known by the nickname "the rat", "SuperRat" or "King Rat" because of his prominent buck teeth.
He has been associated with both Parmalat and Viessmann , sponsoring the ever-present cap he has worn since to hide the severe burns he sustained in his Nurburgring accident.
In the Austrian post office issued a stamp honouring him. Lauda has two sons with first wife, Marlene Knaus married , divorced Lauda has a son, Christoph, through an extra-marital relationship.
In , he married Birgit Wetzinger, who is 30 years his junior and was a flight attendant for his airline. She donated a kidney to Lauda when the kidney he received in a transplant from his brother, in , failed.
On 2 August it was announced that Lauda had successfully undergone a lung transplant operation in his native Austria.
Lauda himself made a cameo appearance at the end of the film. Lauda appears in an episode of Mayday titled " Niki Lauda: Testing the Limits " regarding the events of Lauda Air Flight To Hell And Back: Edit Read in another language Niki Lauda.
Februar bis Laut Lauda erhielt er nur den Geldbetrag von Lauda wurde ehrenhalber in die Lord Jim Loge aufgenommen.
Diese Statistik umfasst alle Teilnahmen des Fahrers an der FormelWeltmeisterschaft , die bis als Automobil-Weltmeisterschaft bezeichnet wurde.
Monaco Monte Carlo Frankreich Le Castellet USA Watkins Glen Ansichten Lesen Bearbeiten Quelltext bearbeiten Versionsgeschichte.
In anderen Projekten Commons. Diese Seite wurde zuletzt am Januar um Weltmeister , , Starts Siege Poles SR Formel 2 Platz 10 ETCC Platz 31 Formel 2 Platz 5 Britische Formel 2 Meister Formel 1 Platz 17 ETCC Platz 7 This is the normal position for the valve without hydraulic pressure applied.
Further examination of the spring that holds the second stage spool in position indicated that it was intact.
The examination of the DCV also revealed that 3 of 4 screws used to secure the solenoid operated pilot valve body to the DCV were loose.
Soil was found inside internal passages of the valve. A metal plug, identified as a case relief valve plug used elsewhere in the engine accessory section, was found installed "finger tight" in the DCV "retract" port.
All solenoid operated pilot valve first stage spool internal passages were unobstructed. There was no evidence that indicated preimpact failure of the valve, however the condition of the valve indicated that the valve was partially disassembled and reassembled by persons not associated with the accident investigation prior to examination by the investigation team.
Additional system tests were performed using production components in an attempt to simulate potential failure modes.
In one hypothetical condition, the introduction of a damaged piece of O-ring seal into a hydraulic orifice resulted in an uncommanded opening of the directional control valve DCV.
For further information on these tests, see paragraph 2. Testing of the electrical function indicated possible areas where an electrical hot short occurring simultaneously with an auto-restow action could result.
A full hydraulic set-up was used to verify normal operation of the thrust reverser system and to determine if uncommanded deployment could occur under various hypothetical failure conditions.
Hypothetical failure conditions involved the directional control valve DCV seal damage, thrust reverser actuator piston head seal leakage and a return line blockage during hydraulic isolation valve HIV cycling.
In another hypothetical failure condition, the effects of piston seal leakage through a thrust reverser actuator was examined with the HIV open.
Several test configurations were examined with the piston head O-ring and cap strip missing from the actuator s. Only one side one of two sleeves of the thrust reverser cowl deployed when an actuator was tested with the piston head seal missing and the bronze plating separated from the piston head.
Under this condition, with the HIV open, internal leakage across the piston was sufficient to deploy the 3 actuators associated with the deployed sleeve depending on the location of the actuator piston head in the cylinders.
If in the stow position and the piston heads were firmly bottomed against the inner cylinder head end prior to commanding thrust reverser stow, the thrust reverser actuators would not deploy.
When the head end of the two actuators were slightly unseated, fluid could pass from the rod end to the head end of the locking actuator causing unlock and extension of 3 actuators one sleeve.
The cap strip from this actuator piston head had considerable wear and was extruded. A DCV was mounted on a vibration table and subjected to resonant searches, resonant dwells, random vibration and sweeps through engine speed.
Pressure transducers and flow meters on the outflow of the valve indicated that the valve did not open unexpectedly or leak during the test under excessive vibration.
The thrust reversers are approved for ground operation only. A general systems description is included in this report as appendix C.
The FAA issued information on the accident to appropriate operators and authorities on September 11, by letter format.
It is included in this report as appendix E. AD , July 3, - Requires tests, inspections and functional checks of the thrust reverser systems on all B airplanes powered by Pratt and Whitney PW series engines.
This superseded AD This superseded TAD 1. AD 9 , October 11, - Requires modification and allowed re-activation of thrust reverser systems on all B airplanes powered by Pratt and Whitney PW series engines.
This superseded TAD Since this information was critical to the investigation, a search was conducted to identify non-volatile memory in various computerized components as an alternate source of data.
The data developed proved helpful in validating conditions prior to and during the accident, but did not provide the time correlation normally available with the DFDR.
The airplane was certificated, equipped and maintained according to regulations and approved procedures. Flight documents indicate that the gross weight and c.
The weather in the area was fair at the time of the accident. Although there were no reported hazardous weather phenomena, isolated lightning was possible.
There are few visible landmarks and population centers on the ground along the route of flight and it is possible that the horizon was not distinguishable.
Recovery from any unusual flight attitude could have been affected by the lack of outside visual references. The pilot-in-command stated "that keeps coming on.
This indication appears when a fault has been detected in the thrust reverser system. It indicates a disagreement. No corrective actions were necessary and none were identified as taken by the crew.
The co-pilot read information from the Airplane Quick Reference Handbook as follows: Airplane design changes implemented after this accident eliminated the need for operational guidance for the flightcrew.
Review of the thrust reverser system design indicates that when the auto-restow system function is required, system pressure to close the reversers is applied during restow and for 5 seconds after restow is sensed.
The specific interval of illumination of the light, and the possibility that the light ceased to be observed, could not be determined from the cockpit voice recorder comments nor from any other evidence.
There was no recoverable data from the nonvolatile memory available in the recovered EICAS components. At ten minutes twenty seven seconds into the flight, the co-pilot advised the pilot-in-command that there was need for, "a little bit of rudder trim to the left.
It ended with the pilot-in-command saying "O. It is probable that the trim requirement was a normal event in the flight profile.
The trim requirement does not appear to be related to the upcoming reverser event, and there was no apparent reason for the crew to interpret it as such.
The physical evidence at the crash site conclusively showed that the left engine thrust reverser was deployed.
Nonvolatile computer memory within the electronic engine control EEC indicated that an anomaly occurred between channel A and B reverser sleeve position signals.
It was concluded that this anomaly was associated with the thrust reverser deployment of one or both sleeves. The EEC data indicated that the thrust reverser deployed in-flight with the engine at climb power; based on EEC design, it was also concluded that the engine thrust was commanded to idle commensurate with the reverser deployment, and that the recorded mach number increased from 0.
The left EEC data indicates that the fuel cutoff switch was probably selected to cutoff within 10 seconds of thrust reverser deployment. Examination of the cutoff switch also indicates that it was in the cutoff position at impact.
A breakup altitude estimation was attempted using time-synchronized information from the CVR. Although the airspeed history between reverser deployment and the end of the recording due to structural breakup cannot be confirmed, the high speeds likely achieved during the descent indicate that the in-flight breakup most likely occurred at an altitude below 10, feet.
Damage to the fan runstrips sic on both engines indicates nontypical loads from an unusual flight path. The fan rubstrips are located on the forward case of each engine and form the fan blade tip airseal.
Each engine fan runstrip sic had a deep rub from the fan blades. The character of the rubs is typical of rubs caused by the interaction with the rotating fan.
The depths are substantially deeper than typical rubs experienced during normal operation. These rubs were centered at approximately 66 degrees on the left engine and approximately 0 degrees on the right engine as view from the rear of the engine looking forward.
Flight testing of the B with JT9D-7R4 engines showed rubs near the top of the engines to be minor depth and centered at approximately 45 degrees on the left engine and approximately degrees on the right engine.
The rub results from aerodynamic load from the engine cowls. These loads were determined to be essentially down from the top when the aircraft nose was lowered during descent.
The PW installation is designed for the maximum cowl aerodynamic loads that occur during takeoff rotation. At that condition a.
This rub would be due to upward aerodynamic force on the cowl at aircraft rotation angles of attack. The depth and location of the rubs in the.
Lauda accident indicates; 1 cowl load forces much greater than the forces expected during takeoff rotation and 2 by the location, that the forces were essentially down from the top of the cowl.
The CVR transcript indicates that the in-flight breakup did not occur immediately after the deployment of the thrust reverser, but rather during the subsequent high-speed descent.
The EEC can provide general altitude and Mach number data however calibration is not provided outside the normal speed envelope. Information from the engine manufacturer indicates that the EEC data may indicate altitude and Mach numbers which are higher than the true value.
Also, EEC calibration of its ambient pressure sensor affects the accuracy of the recorded Mach number and altitude. This calibration is not designed to be accurate above maximum certified airplane speeds.
In addition, the EEC ambient pressure calibration does not account for the effect of reverse thrust on fan cowl static pressure ports. However, EEC recorded data does suggest that the airplane was operating beyond the dive velocity of 0.
High structural loading most probably resulted as the crew attempted to arrest the descent. Parts of the airplane that separated from buffeting overload appear to be pieces of the rudder and the left elevator.
This was followed by the down-and- aft separation. No evidence of impacts were observed on the leading edges of the horizontal and vertical stabilizers indicating that no airframe structural failure occurred prior to horizontal stabilizer separation.
It is thought that the download still present on the left stabilizer and the imbalance in the empennage from the loss of the right stabilizer introduced counterclockwise aft looking forward orientation torsional overload into the tail, as evidenced by wrinkles that remained visible in the stabilizer center section rear spar.
The separation of the vertical and left horizontal stabilizers then occurred, although the evidence was inconclusive as to whether the vertical stabilizer separated prior to or because of the separation of the left stabilizer and center section.
The damage indicated that the vertical stabilizer and the attached upper portion of four fuselage frames departed to the left and that separation of the vertical fin-tip and the dual-sided stringer buckling in the area of the fin-tip failure occurred from bending in both directions prior to the separation of the vertical stabilizer from the fuselage.
The loss of the tail of an airplane results in a sharp nose-over of the airplane which produces excessive negative loading of the wing.
Evidence was present of downward wing failure. This sequence was probably followed by the breakup of the fuselage.
The complete breakup of the tail, wing, and fuselage occurred in a matter of seconds. 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. 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.
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 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 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.