TM 1-1520-238-10 2-26 Change 9 NOTE The ECU and DECU are not inter- changeable between -701 and -701C en- gines. 2.26.5 Electrical Control Unit (ECU) . The ECU (fig 2-16) controls the engine and transmits operational in- formation to the crew stations. It is a solid-state device mounted below the engine compressor casing. Powered by the engine alternator, the ECU receives inputs from the thermocouple harness, N_p sensor, torque and overspeed sensor, opposite engine torque for load sharing, N_G signal from the alternator, N_p reference signal from the turbine speed control unit, and a feedback signal from the HMU for system stabilization. The torque-sharing system in- creases power on the lower-output engine to match it with the higher output engine. The ECU also receives opposite engine torque inputs to enable contingency power. When this input signal is 51% torque or below, contingency pow- er is automatically enabled. However, contingency power is not applied until the flight crew pulls in collective above 867 C TGT. The ECU automatically allows the normally operating engine to increase its TGT limit, thereby in- creasing its torque output. The overspeed protection sys- tem senses a separate N_p signal independently of the governing channel. ECU also provides signals to the N_p indicator, TORQUE indicator, and history recorder. In case of the ECU malfunction, system operation may be overrid- den by momentarily advancing the engine PWR lever to LOCKOUT and then retarding the lever past the FLY posi- tion to manually control engine power. This locks out the ECU from all control/limiting functions except N_p over- speed protection, which remains operational. To remove the ECU from lockout operation, the engine PWR lever must be moved to IDLE, then back to FLY. a. Engine TGT Limiter Function . The ECU in- corporates a steady state dual and single engine TGT lim- iting function which restricts fuel flow within the HMU to prevent an engine over–temperature. The limiting function has an inherent 4 C variance. The resistance in the cab- ling and circuitry between the ECU and TGT gauge is enough to produce a 5 C variance factor. Applying the sum of these two factors, the dual engine limiter setting is allowed a value of 860 9 C (851 – 869 C) and the single engine (contingency power) limiter setting is al- lowed a value of 917 9 C (908 – 926 C). The TGT lim- iter setting for a particular engine can change within these ranges over a period of time. 2.26.6 Digital Electronic Control Unit (DECU) . The DECU (fig 2-16) is mounted in the same loca- tion as the ECU. The DECU can be overridden like the ECU by momentarily advancing the engine PWR lever to LOCKOUT. The DECU, which incorporates improved technology, performs the same functions as the ECU ex- cept for the following functional and control improve- ments. The DECU can be fully powered by either the engine alter- nator or by 400 Hz, 115 vac aircraft power. It incorporates logic which will eliminate torque spike signals during en- gine start-up and shutdown. The DECU control logic con- tains a Maximum Torque Rate Attenuator (MTRA) feature designed to reduce the risk of exceeding the dual engine torque limit during uncompensated maneuvers. These are any maneuvers where pedal or cyclic inputs are made but no collective control inputs occur. For example, large tran- sient engine torque increases can occur during left pedal and left lateral cyclic inputs when performing maneuvers such as rapid hovering turns or forward flight roll rever- sals. The MTRA is designed to reduce fuel flow and limit rate of torque increase to approximately 12% per second when at transient engine torque during an uncompen- sated maneuver. However, any collective control increase during the maneuver can override the MTRA and normal maximum engine torque rate increase can be achieved. The MTRA feature is not active in the DECU logic under single engine conditions. The DECU contains an automat- ic hot start preventer (HSP). The DECU also provides sig- nal validation for selected input signals within the electri- cal control system. Signals are continuously validated when the engine is operating at flight idle and above. If a failure has occurred on a selected input signal, the failed component or related circuit will be identified by a pre-se- lected fault code. Fault codes will be displayed on the en- gine torque meter (fig 2-18), which defines fault codes in terms of engine torque. Fault codes will be displayed start- ing with the lowest code for four seconds on/two seconds off, rotating through all codes and then repeating the cycle. The fault codes will be displayed on the engine torque meter only when all of the following conditions are met: N_G less than 20% N_p less than 35% Other engine shutdown Aircraft 400 Hz power available The fault codes can be suppressed by pressing either OVSP TEST switch. The fault codes can be recalled by again pressing either OVSP TEST switch. Once a failure has been identified, the fault code will remain available for diagnostic indication until starter dropout on the next en- gine start.