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Under NASA contract, PCA system preliminary design and development studies were conducted in 1993, with prime contractor MDA establishing subcontracts with Honeywell (manufacturer of .ight control computer and software) and engine contractors, Pratt & Whitney and General Electric (Evendale, Ohio). The preliminary design was completed in the fall of 1994. The study showed that an all-software PCA implementation was feasible, and performance appeared to be adequate for the project objectives. A test airplane with Pratt & Whitney engines was selected for the demonstrations.
Final design subcontracts were let to Honeywell and Pratt & Whitney, and the .nal design was completed in March 1995. Pratt & Whitney modi.ed a breadboard FADEC, tested the modi.ed software on a test engine at sea-level static conditions, and then programmed read-only memory modules for four sets of .ight-quali.ed FADECs. Honeywell developed the software for the PCA control laws, with gains based on the control evaluations conducted in the FDS. This PCA software was tested in an of.ine simulation at the Honeywell facility. After this software was operating properly, it was coded for the FCC; unneeded software modules were deleted as required to make room for the PCA code.
The .ight code was then tested by Honeywell. MDA conducted integration tests in the MD-11 closed-loop bench simulation with a PCA FCC and a single test PCA FADEC to verify proper communication over the data buses as well as proper engagement, annunciation, operation, and disengagement (ref. 14). The MD-11 bench setup could not accommodate more than one FADEC. A ground test on the airplane veri.ed end-to-end operation, including that PCA commands could be sent to the FADECs, that the engines would respond properly, and that the PCA system could be disengaged with all methods.
Test Procedure
For PCA system .ight tests, the MD-11 test airplane 560 was con.gured as usual; LSAS and yaw dampers were turned off and speed-protect logic was inhibited by lowering the FCC paddle switch. The left-seat pilot was the PCA test pilot; the right-seat pilot was the safety pilot and did not participate in PCA system control. Table 1 shows the test pilots. The .ight test engineer sat behind the center console, set up the test conditions with the pilots, veri.ed aircraft con.guration, made MCDU entries, and recorded .ight notes. In the main cabin, the data system engineer veri.ed data acquisition and coordinated with test engineers viewing data in real time at display consoles. The Flaps 28 (maximum takeoff .ap) setting was used for much of the PCA system .ight testing, as this would be the case right after takeoff and on approach. Also, most of the PCA system testing was performed with the gear down. This con.guration was used because, in an emergency, the gear could be lowered without hydraulic pressure, and the increased drag from the gear was bene.cial for PCA system operation. Tests were .own from the MDA test facility at Yuma, Arizona.
Table 1. PCA system pilots.
Pilot Af.liation Title Last name
A NASA NASA project pilot Fullerton
B MDA MD-11 PCA project pilot Luczak
C MDA MDA evaluation pilot, MD-11 chief pilot Miller
D NASA NASA evaluation pilot Purifoy
E FAA Demonstration pilot, national resource specialist Imrich
PCA System Flight Tests
The PCA system .ight tests consisted of 21 .ights .own over a 5-month period (table 2). The PCA system .ights were interspersed with .ights for other purposes. The .rst PCA .ight was with a FADEC installed on only one engine for safety reasons. The test software had a switch activated by an MCDU input that permitted operation with a single FADEC. The .rst .ight (.ight 200) was .own in August 1995. Because of a software loading error, no data were obtained. On .ight 201, successful PCA engagement and disengagement tests were made, followed by some small PCA commands to verify correct engine response. Then, with the PCA system off, some open-loop throttle response tests were performed. Figures 7 and 8, shown earlier, demonstrate these open-loop throttle step results with gear down and .aps up. The aircraft response to an increasing throttle step with gear down and Flaps 28, is shown in .gure 15(a), while response to a throttle decrease with gear down but .aps up is shown in .gure 15(b). These data were used in further comparisons with the FDS data.
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