Cincotta v. United States

• Decision Date: 18 June 1973
• Docket Number: 71-1358.,Civ. A. No. 71-1179
• Citation: 362 F. Supp. 386
• Parties: Ruth Marie CINCOTTA, surviving widow of Eugene Joseph Cincotta, Deceased, to her own use and to the use of Continental Casualty Company v. UNITED STATES of America. Carmen S. TURNER, surviving widow of Robert Stanton Turner, Deceased, to her own use and to the use of Continental Casualty Company, Robert Stanton Turner, Jr., a minor, by Carmen S. Turner, his mother and next friend, to his own use and to the use of Continental Casualty Company and Carmen S. Turner, Executrix of the Estate of Robert Stanton Turner, Deceased v. UNITED STATES of America.
• Court: U.S. District Court — District of Maryland

362 F. Supp. 386

Ruth Marie CINCOTTA, surviving widow of Eugene Joseph Cincotta, Deceased, to her own use and to the use of Continental Casualty Company v. UNITED STATES of America.

Carmen S. TURNER, surviving widow of Robert Stanton Turner, Deceased, to her own use and to the use of Continental Casualty Company, Robert Stanton Turner, Jr., a minor, by Carmen S. Turner, his mother and next friend, to his own use and to the use of Continental Casualty Company and Carmen S. Turner, Executrix of the Estate of Robert Stanton Turner, Deceased v. UNITED STATES of America.

Civ. A. Nos. 71-1179, 71-1358.

United States District Court, D. Maryland.

June 18, 1973.

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Patrick A. O'Doherty, Baltimore, Md., for plaintiff Ruth Marie Cincotta.

Thomas E. Lloyd and Robert F. Fischer, Ellicott City, Md., for plaintiffs Carmen S. Turner and Robert Stanton Turner Jr.

George Beall, U. S. Atty., D. of Maryland, and George E. Farrell and Joseph T. Cook, Dept. of Justice, Washington, D. C., for defendant in both cases.

NORTHROP, Chief Judge.

This is a wrongful death action against the United States of America brought pursuant to the Federal Tort Claims Act, 28 U.S.C. §§ 1346(b), 1402 and 2671 et seq. (1970). The plaintiffs are Ruth Marie Cincotta, widow of Eugene Joseph Cincotta, and Carmen S. Turner, widow of Robert Stanton Turner. The factual issues are of a very technical nature, and thus extensive background material must be initially discussed.

GENERAL AND TECHNICAL BACKGROUNDS

On December 16, 1969, a flight test of a B-57G airplane was scheduled at the Martin-Marietta Corporation, Baltimore Division, as part of a test program for the United States Air Force. The airplane was piloted by Robert Stanton Turner and Eugene Joseph Cincotta flew in the rear seat as a test observer. One of the tests to be flown was a demonstration of Minimum Single Engine Control Speed (VMC). During the execution of this exercise, the airplane went out of control, became inverted and crashed into the Sassafras River in Cecil County, Maryland. Both Mr. Turner and Mr. Cincotta were instantly killed.

The Crew

Mr. Turner was an experienced and competent test pilot. He was Manager and Chief Test Pilot of the Flight Test Department of the Martin-Marietta Corporation, Baltimore Division (hereinafter referred to as MMC). Mr. Turner had a total flying time of 4,086 hours, with 1,247 hours of that being in jet aircraft and 715 hours in B-57 airplanes. A 1953 graduate of the United States Air Force Experimental Test Pilot School, Mr. Turner had amassed a total of sixteen years' flight test experience. He was considered an outstanding and highly competent test pilot by his fellow test pilots.

Mr. Cincotta was employed by MMC in the capacity of a flight test engineer. Although not a test pilot, he was an experienced and qualified pilot and was actively flying for the Maryland Air National Guard. In connection with the B-57G tests, Mr. Cincotta flew in the rear seat of the aircraft as a test observer. His duties included operation of instrument switches, monitoring test conditions, knee board recording of test parameters and assisting the test pilot as required. It is not disputed that he was well-qualified for this position of test observer.

The Aircraft

The aircraft involved in the fatal accident at issue was a B-57G, which is a modification of the B-57B,¹ an aircraft originally manufactured by the Martin Company as a light bomber. The B-57B was the first American modification of a British designed airplane acquired by the United States Air Force — the B-57. The airplane is considered extremely versatile and can be used for a variety of missions including electronic counter-measures, air sampling, research and development, and other special purposes. Through the years the original B-57B has undergone further modifications into various different aircraft series — the B-57C and the B-57E. The particular airplane involved in this case was originally accepted by the Air Force as a B-57B on October 4, 1955, and was given a serial number of 53-3905. The aircraft was subsequently returned to MMC under a bailment agreement² for modification into a B-57G and related testing. The modification was completed on September 5, 1969, at the Baltimore Division of MMC. Prior to the fatal flight the aircraft, as modified, had flown for 59.4 hours during the conduct of some 31 tests related to the modification program.

The only component of the aircraft requiring extended description is the rudder control system, for that is the only portion of the aircraft involved in the instant litigation. Despite this limitation, the technical discussion required remains extensive, because the rudder control system is complex. Nevertheless only a rudimentary explanation is possible. It is hoped, however, that the attached diagrams will aid in the understanding of the discussion.

The rudder control system³ consists of combination rudder and brake pedals, a jackshaft, push-pull rods from the jackshaft to the rudder, a torque tube and blow-back rod assembly in the rudder leading edge, a rudder, and a combination spring and trim tab. The rudder pedals are connected by a torque tube to a lever below the pilot's floor. Attached to this lever is a push-pull rod, the opposite end of which is attached to another lever on a jackshaft. The jackshaft is left of the airplane center line and has an upper and lower lever mounted on it. The lower lever attaches to a push-pull rod system which runs aft through a pressure box to the torque tube and blow back rod assembly in the rudder leading edge. Thus, fore and aft motion of the rudder pedals is transmitted aft through the push-pull rod system to the slotted lever on the left side of the torque tube and blow-back rod assembly in the rudder leading edge. Movement of the slotted lever turns the rudder through the torque tube. This system is depicted clearly in Figure 3.⁴

The important feature of the Rudder Control System, at least as far as this case is concerned, is the hydraulic rudder power assist system which assists the pilot in moving the rudder beyond certain limits of rudder pedal or trim displacement. Figure 4⁵ depicts the Rudder Power Assist System which consists of a push rod assembly, an actuator assembly, and a cylinder pick-up arm. The push rod assembly is attached at one end to the top part of the rudder system bell crank, forward of frame 42, and at the other end to the actuator assembly valve lever aft of frame 42. The actuator assembly is a combination actuator and valve enclosed in the same housing. The actuator assembly valve is operated mechanically by the valve lever attached to the push rod assembly. The rod end of the actuator is attached to a bracket on frame 42 and the other end is attached to the cylinder pick-up arm. The cylinder pick-up arm is splined and mounted on a splined shaft attached to the bottom rudder hinge bracket.

The push-rod assembly and the actuator valve lever move with the normal rudder system linkage when the system is not in operation in normal flight. (This is because the slide valve linkage has a wide dead band around zero.) Whenever the rudder trim tab is operated more than fourteen degrees in either direction, fingers on the valve lever contact a washer on the end of the value spool. Movement of the value spool opens and closes ports to direct hydraulic fluid into the actuator under pressure. Pressure into either side of the actuator moves the actuator, the cylinder pick-up arm, and the rudder. During normal flight the rudder trim tab is never displaced more than 14°, and thus the actuator does not come into play. The system is, in fact, only required in low speed single engine conditions.

As trim tab deflection approaches fourteen degrees, rudder pedal force necessary to activate power assist becomes less. The power output from the actuating cylinder is applied directly to the rudder and results in greater deflection than is usually available. The point at which rudder assist is initiated is not apparent to the pilot due to the air loads acting on the surface of the rudder.

The Test

On December 16, 1969, the aircraft flown by Mr. Turner was participating in test flight number 35 which was intended to obtain both Static Longitudinal Stability data and VMC (Minimum Single Engine Control Speed) data. At the time of the fatal crash the crew was engaged in a maneuver which constituted an attempt to demonstrate the Minimum Single Engine Control Speed.

A VMC test is conducted in the takeoff configuration, which means that the landing gear are to be down and the wing flaps are to be up. There are to be no external stores and the aircraft must be trimmed for takeoff configuration with the power assist on and the gross weight at the lightest normal takeoff loading. This test is conducted at the minimum safe altitude.⁶ Further, one engine must be failed⁷ and the other operating at takeoff power, the airspeed being reduced while maintaining altitude or slight climb. Straight flight must be achieved and maintained with no greater than a 5 degree bank, 180 pounds rudder force or full rudder deflection, whichever comes first. The speed at which either of these two criteria are exceeded is the minimum single engine control speed and must be equal to or less than 1.2 VSTO. (1.2 of the stall speed for takeoff configuration.)

Two months prior to flight 35, Mr. Turner was successful in stabilizing the airplane in straight flight at 134 knots, a figure which was within the test requirements for meeting minimum single engine control speed. This was accomplished during flight 11 on October 3, 1969. Initially, the United States Air Force refused to accept the results of flight 11 due to certain difficulty with the synchronization of the flight data recording system, and a reflight, test flight 35, was called for. Subsequently, the Air Force changed its...