On December 13, 2001, at 1203 central standard time, a Boeing 737-322, N359UA, operated by United Airlines (UAL) as flight 578, declared an emergency due to a reported "rudder malfunction" during decent to the Chicago O'Hare International Airport (ORD), Chicago, Illinois. Visual meteorological conditions prevailed at the time of the incident. The 14 CFR Part 121 domestic passenger flight was operating on an instrument rules flight plan. The 2 pilots, 4 flight attendants, and 93 passengers reported no injuries. The flight departed from the Lambert-St. Louis International Airport, St. Louis, Missouri, at 1129, en route to ORD.

The first officer stated that while descending through 9,300 feet mean sea level (MSL) the airplane began to bounce and slowly bank to the left, which led her to believe that they were experiencing wake turbulence. The A autopilot was engaged at the time when the airplane began a correction back to the right while continuing to bounce. The airplane rapidly banked 10 degrees when the captain said that he was disengaging the autopilot and autothrottles. The first officer described the acceleration rate of the bank as when the airplane is turned with the speed brakes deployed to the flight detent. Due to the yawing and bouncing sensation, the first officer instructed the flight attendants via public address system to be seated immediately. The first officer reported the airplane was now in about a 45-degree right bank.

The first officer further stated that the captain was very obviously opposing the roll with left rudder pressure and left aileron. The captain said something to the effect, 'I think it's a rudder problem.' The first officer stated that the A and B system hydraulic quantity and pressure gauges indications were normal. The first officer began assisting the captain by applying left aileron and left rudder control pressure and noticed that the airspeed decreased from 250 knots indicated airspeed (KIAS) to 238 KIAS. In response, the first officer pushed forward on the control column to increase the airspeed in an attempts to increase control effectiveness. As the airspeed increased, the effect of left control inputs began to move the airplane to a more comfortable attitude.

The captain appeared to regain control of the airplane about 8,200 feet MSL. At this point, the captain still had both hands on the control wheel with significant left aileron input. The first officer increased engine power to maintain 250 KIAS and 8,000 feet MSL. The captain asked the first officer to declare an emergency. The first officer reported a rudder malfunction, 100 souls on board, and about 10,300 pounds of fuel. The first officer initiated the quick reference checklist (QRC) for an uncommanded rudder, but the reference actions in the QRC did not provide any relief to the amount of the left aileron and left rudder pressure required to maintain control of the airplane. The first officer stated that there was no change in hydraulic pressure and quantity indication upon completion of each checklist item.

ORD Approach Control cleared flight 578 to 4,000 feet MSL and issued vectors for the runway 27L downwind. They were then cleared to descend to 2,500 feet MSL and 210 KIAS. At 220 KIAS, the captain called for flaps 2 which was then selected by the first officer. The captain then said that it "looks like a slam dunk." The captain utilized the speed brakes and called flaps 5. Air traffic control turned the flight on a base leg and cleared it to 2,100 feet MSL and 180 KIAS. The captain stowed the speed brakes. About 195 KIAS, the captain called for flaps 10 and gear down final descent checklist. At this time, they were issued a final vector and cleared for the ILS 27L approach. They flew a flaps 30 approach and landing to runway 27L. The first officer stated that the captain was stabilized per UAL criteria well before 1,000 feet above ground level (AGL). The captain commented, between 500 feet and 1,000 feet AGL, how it was becoming increasingly difficult to control the airplane with left rudder pressure and left aileron. At 500 feet AGL, the first officer assisted the captain on the control per his request.

The airplane landed on the centerline and touchdown zone of runway 27L. The speed brakes automatically deployed upon wheel spin up. The captain used some reverse thrust together with the autobrakes, began to slow the airplane. The captain and first officer attempted to maintain the directional control of the airplane through the landing roll with the use of rudder input, differential braking, and asymmetrical thrust. The first officer stated that it became increasingly difficult as the airplane slowed. About 100 KIAS, the first officer heard a loud ping or thud when the rudder simultaneously locked out all control input causing the left pedal to push against the first officer's foot.

The captain stated that during landing roll, about 100 KIAS, they heard a banging noise and got a full right scale deflection of the rudder pedals. The rudder pedals freed up and seemed back to normal shortly after coming to a stop and setting the parking brake. The airplane was then towed to the gate.


Federal Aviation Regulation 121.359(h) Cockpit Voice Recorders, states:

"In the event of an accident or occurrence requiring immediate notification of the National Transportation Safety Board under part 830 of its regulations, which results in the termination of the flight, the certificate holder shall keep the recorded information for at least 60 days or, if requested by the Administrator or the Board, for a longer period. Information obtained from the record is used to assist in determining the cause of accident or occurrences in connection with investigations under part 830. The Administrator does not use the record in any civil penalty or certificate action."

The cockpit voice recording of the flight was not retained. The CVR and digital flight data recorder (DFDR) circuit breakers were not pulled out by either the flight crew or maintenance personnel following the landing. Data from the DFDR relating to the incident flight was able to be retrieved by the Safety Board.

A second DFDR was installed to record a series of control sweeps performed by the Airworthiness Group during the investigation.

The DFDR recording of flight 578 shows that the airplane achieved a right 26-degree roll; a left 65-degree control wheel deflection followed by right control wheel deflections. The rudder pedal deflection achieved a maximum value from left 10 degrees to right 10 degrees, which was followed by a right 1.5 degree deflection with corresponding aileron deflections ranging from left 6.0 degrees to right 1.0 degree. The pedal then remained deflected to the left over a range from 2.0-0.25 degrees with corresponding left rudder deflections ranging from 2.5-0.25 degrees until the airplane reached an altitude equal to about ORD airport elevation at which point both rudder and elevator deflections began to occur to the left and right center.


Flight 578 was a designated UAL Aircraft Communication Addressing Reporting System weather flight. At 1146, the maximum pressure altitude recorded during the flight was 26,990 feet with an outside air temperature (OAT) of -35.8 degrees C. At 1200, the maximum eddy dissipation rate was recorded as 0.35, which equates to moderate turbulence. An OAT of -3.3 degrees Celsius (C) was also recorded at this time.

The ORD Automated Weather Observing System, recorded at 1156: wind 250 degrees at 8 knots; visibility 10 statute miles; sky condition 1,500 feet overcast; temperature 6 degrees C; dew point 3 degrees C; altimeter setting 29.93 inches of mercury.


The Airworthiness Group performed visual inspections and testing of the rudder and lateral control system while following, but not limited to, guidance provided in service letter 737-SL-27-110-A, Unexpected Roll and Yaw Event Troubleshooting. Visual inspections of the flight control system, functional inspections, and testing were performed before any component removal.

Cockpit inspection revealed that the captain's left rudder pedal and first officer's right rudder pedal made intermittent contact with plastic housing located between the pedals. Nuts, drilled rivets, and other foreign objects were found in the foot wells. The housings were removed exposing the underlying cockpit floor and a portion of the rudder pedal arms. The cockpit floor aft of the pedals, the surrounding structure, and pedal arms were covered with lint/dirt ranging in thickness upwards of 1/2 inch. About 60 percent of the underlying flooring possessed a continuous area of reduced lint/dirt.

Inspection of the fuselage nose area forward of the nose landing gear revealed a wood-handled brush with wire bristles inventoried by UAL as Brush-Wire 1-1/2" 3 Row 8"L Stainless Steel, Part Number BR770. The brush, along with a fragment of twisted safety wire, drilled rivets, a nut, and plastic retainers were beneath the first officer's rudder pedals. The first officer's pedal arms contained scuff marks on the paint surface and gouge marks which penetrated the paint surface and metal. Further inspection with the aid of a borescope revealed the presence of displaced white paint on the structure aft of the pedal arms. No dents were noted in the vertical structure surrounding the pedal arm. The UAL party representative stated that work relating to an airworthiness directive was performed weeks prior to the incident. The brush, along with a new brush and the first officer's rudder pedal arms, were sent to the National Transportation Safety Board's Materials Laboratory for further examination.

According to the Materials Laboratory Factual Report, which is included in the docket of this report, the examination of the left rudder pedal arm in the vicinity of the bolt heads revealed only light score marks scattered around the edge of the boss, consistent with tool usage. The right rudder pedal arm contained one score mark on the rear edge of the boss. The score mark was circumferential for approximately 90 degrees of arc, centered at the horizontal line. It was more significant above the horizontal and tapered to nothing below it.

The incident brush consisted of a plywood handle with 13 groups of metal wires at one end. The plywood handle consisted of seven veneers, four with a nominal thickness of 0.06 inches and three with a nominal thickness of 0.04 inches. Examination of the left, right, top, and bottom faces of the handle revealed no indications of any crush damage.

Blue stains were visible beneath the lavatory area did not extend onto flight control cables. Electrical connectors behind the yaw dampener were absent of stains and wetness. The cables in this area had up to a 1/2 inch of lint/dirt on them. The cables were lifted and the pulley were then rotated by hand; no binding or jamming was noted.

The mid-cabin seats and flooring was removed exposing the flight control cables in the wing center section. The area over the fuel tank was predominately free of foreign objects except for several chips of debris. The cables were lifted and the pulley were then rotated by hand; no binding or jamming was noted.

The overhead liners in the rear cargo compartment were removed exposing the rudder cables. The cables in this area had up to a 1/2 inch of lint/dirt on them. The area beneath the aft lavatories contained blue colored stains, none of which were near the rudder cables.

Inspection of the pressure seals and flight control cable pulleys aft of the rear pressure bulkhead revealed no anomalies. The cables were under tension and no water was found beneath the control paths.

The rudder power control unit (PCU) access panels on the vertical stabilizer were removed. Upon removal, the presence of condensation was reported on the area components. Water and hydraulic fluid was found pooled beneath the PCU. The PCU had white corrosion deposits on the lower left case and fluid stains on the structure leading away from beneath the PCU seals.

A Non-Routine Maintenance Card from the last C-check indicated that fluid had been found, the area cleaned, and no leakage was observed during a pressure check. The PCU emitted a hissing sound when hydraulic power (A, B, and Standby) was applied and no external leakage was observed.

Paint was missing from the aft rudder control torque tube where it connected into the lower bell-crank fitting. At full right rudder trim, the lower standby PCU hose contacted the torque tube.

When the A and B hydraulic systems were pressurized, the standby PCU was silent after it emitted a single "knock" sound during initial movement. Visual inspection of the area noted a single loose washer on the fixed structure beneath the movable trim frame. The frame for the feel-trim-centering unit was 0.010 inches from a structural rivet beneath the movable portion of the unit. Inspection of the PCU from the right side of the tail revealed curvilinear scrapes were present on the structure underlying the PCU and the top half of a rivet head was missing. The PCU was actuated with A and B hydraulic systems pressurized via the rudder pedals. The PCU followed the scrape path without contacting the surface even with downward hand pressure. A Boeing engineer noted that the scrapes on the structure could have been the result of installation of the PCU.

The trim motor wires were found stretched and no slack existed. The Boeing Standard Practices Wiring Manual calls for a small amount of slack in wire bundles. A subsequent operational check revealed the trim motor was functional.

A Radar Audio Playback Terminal Operations Recording (RAPTOR) was created by the FAA and is included in the docket of this report. The RAPTOR shows flight 578 preceding UAL flight 108, heavy.

The captain's and first officer's rudder breakout forces were 8-9 pounds. The Boeing specification cites a maximum breakout force of 18 pounds.

The QRC Uncommanded Rudder checklist lists the following:

Maintain control of the airplane with all available flight controls. If roll is uncontrollable, immediately reduce pitch/angle of attack and increase airspeed. Do not attempt to maintain altitude until control is recovered.

Verify thrust is symmetrical.

Reference Action

The Uncommanded Rudder procedure should be followed for either rudder pedal displacement or pedal kicks.

Yaw damper switch............................................Off
Do not use Yaw Damper Inoperative Irregular procedure.

Rudder trim........................................................... Center
Rudder pedals..................................................... Free and center
Use maximum force, included a combined effort of both pilots, if required, to free and center the rudder pedals.

If rudder pedal position or movement is not normal and the condition is not the result of rudder trim:
System B flight control switch............................Standby rudder


Use flaps 30 or 40 for landing (provides increased roll control due to increased spoiler effectiveness).
A slight rudder deflection may remain, but continued rudder pedal pressure may help maintain an in-trim condition.
Sufficient directional control is available on landing using differential braking and nosewheel steering.
Crosswind capability may be reduced
Do not use autobrakes
Consider checking rudder freedom of movement at a safe altitude using slow rudder inputs while in the landing configuring and at approach speed.

If the condition was the result of rudder trim or environment factors:

Yaw damper switch.........................................................................................On
Normal Approach Descent and Final Descent Checklist....................Accomplish

According to "Dynamics of Flight Stability and Control" (1996), Estimates of the Maximum Rudder Forces that Can be Exerted for Various Positions of the Rudder Pedal (BuAer, 1954) lists the following values:

Rudder Pedal Position Distance from Back of Seat (in) Pedal Force (lb)

Back 31.00 246
Neutral 34.75 424
Forward 38.50 334

Examination and testing of the brush revealed the following. The brush was an 8-inch plywood-handle wire brush. Several brushes of the same type were loaded in the lengthwise direction and also placed various obstructive positions within the rudder pedal arms. Brushes were placed in obstructive positions with a pilot applying rudder pressure. The pilot estimated that the force he applied to the rudder pedals to fracture the brushes was 15-20 pounds. Brushes loaded in the lengthwise directions broke at loadings from a maximum of 80 pounds and below.

A flight training simulator was back driven at UAL's Flight Training Facility using FDR data from the incident flight with both flight crew members aboard in order to simulate the incident. The simulator structure at the rudder pedals had the same platform/shelf structure as the incident airplane. The rudder pedals were set to the same adjustment as reported by the first officer, and a brush was inserted with the application of left rudder pedal pressure to hold the brush in place due to the resting platform being shorter than the incident airplane. The brush position for the best strength and holding power was noted when the brush was orientated such that the wire bristles were facing downward and towards the rudder pedal control arms and the convex portion of the brush handle facing upwards with the other end against the vertical wall of the pedal box. Further pressure on the rudder pedals was not added due to a concern of causing damage to the simulator.

The left and right rudder pedal arms from the first officer's rudder controls, the incident wire brush, and a new sample wire brush were examined at the National Transportation Safety Board's Materials Laboratory. The Safety Board's Material Laboratory Report is included in the public docket of this report. Examination of the left, right, top, bottom, and rear faces of the incident brush handle revealed no indication of any crush damage. The UAL party representative reported that there had been recent maintenance work, which required the use of tools in the area near the pedal controls.

The PCU, standby PCU, and both flight control computers (FCCs) were removed and shipped to their respective manufactures for testing. The PCUs were shipped to Parker Aerospace, Irvine, California, and the flight control computers were shipped to Honeywell, Renton, Washington.

PCU testing was within the parameters described in Boeing Overhaul Manual (OHM) 27-20-01 except for the Coil Resistance Check, which called for a resistance between solenoid pins one and two to be 72-76 ohms; the measured resistance was 77 ohms. The following tests called for in the OHM were not performed: Continuity Check, Dielectric Strength, Return Pressure, Proof Pressure, Cylinder Rod Leakage, Internal Leakage, Intersystem Leakage, Duty Cycle, and Low Pressure Leakage.

The PCU was disassembled and the dimensions of the primary summing lever and the primary slide (servo) were measured. The outer diameter of the lever ball was 0.3123 inches. The diameter of the mating bore on the slide was 0.3126 inches. Both diameters and their clearances were reported to be within the wear limits specified in the OHM. A macroscopic visual inspection of the slide's right and left lands revealed scratches/wear, which was described by Boeing as not having excessive scratches, wear, or other anomalies.

The servo valve was removed and tested in accordance with acceptance test procedure (ATP) 398310T for the following tests: Secondary Slide Friction and Detent Spring Force, Primary Flow Force, Primary Slide Friction, and Residual Pressure. The servo valve was within the test parameter except for the Secondary Slide Friction and Detent Spring Force. The force required to begin movement of the slide from detent to extend and back were 6.5 pounds and 7.0 pounds, respectively; the requirement is less than 6.25 pounds for both directions. The force difference required to move the slide from detent to retract and to begin movement toward detent from retract was 4.0 pounds; the requirement is less than 1.5 pounds. There are no in-service limits at the component level.

Maintenance monitor information (BITE) was extracted from both FCCs and a full acceptance test procedure run was performed on both units. BITE did not contain any information related to the incident and testing of both units was accomplished within the test's parameters.

A search of the Federal Aviation Administration’s service difficulty reporting system relating to "Boeing Rudder Controls" was performed for a time period from January 1, 1997 to February 1, 2002. One occurrence showed that maintenance could not find any fault with the airplane and that the rudder hydraulic system components were replaced. A second occurrence showed that after the accomplishment of engineering callout relating to rudder hydraulic system components, no defects were noted. A third occurrence noted the replacement of the PCU. These occurrences included the following:

On February 23, 1999, a Boeing 737-2B7, began to pitch up and roll 80 degrees to the right at flight level 330 and 0.74 Mach. The airplane was on approach standby with the 'A' autopilot. The captain stated that when he turned off the 'B' hydraulic pump, the airplane snapped back to normal. The captain declared an emergency and diverted the flight. The flight landed without further incident. Maintenance could not find any fault with the airplane. The PCU, standby PCU, standby hydraulic pump, yaw dampener coupler, number 6 and 7 flight spoiler actuators, and all hydraulic filters were replaced.

On September 15, 1999, a Boeing 737-522, experienced a possible uncommanded rudder deflection during visual approach. The airplane was on a visual approach for landing and 3-4 miles behind a Boeing 737. At the final approach fix, the airplane encountered momentary wake turbulence and then smooth air. The airplane then rolled right. The first officer applied left aileron and the airplane began to respond and then began to roll further right to about 30 degrees. First officer applied full left rudder. The airplane did not respond for several seconds. The captain called for a go-around and the airplane responded. The airplane returned and landed without further incident. The first officer stated that he felt an right rudder travel 1/4 to 1/2 full deflection, uncommanded. No defects were noted upon accomplishment of engineering callouts, a PCU servo valve test, and yaw dampener troubleshooting.

On November 4, 2001, a Boeing 737-490, had a left rudder restriction in flight. The left rudder could be moved about 2 inches and stopped. The standby rudder had more authority but had restrictions. The PCU was replaced and the operational check was normal.


The FAA, Boeing, UAL, and the Airline Pilots Association, International were parties to the investigation.

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