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.