Abstract: When helicopters are accomplishing certain special maneuvers such as powered descent, or the like involving low "g" conditions, the helicopter blades may teeter excessively with respect to the mast, and under severe conditions, the mast may shear. To sense such low "g" conditions, which may be dangerous, an accelerometer is mounted on the helicopter frame close to the center of gravity of the helicopter and is coupled to data processing circuitry which classifies the low-g conditions in bands or levels of severity, recording the duration of exceedances in which the low-g conditions are below predetermined levels, and issues warning signals to the helicopter pilot through the helicopter intercom system and a suitable warning light.

Abstract: A control system for a jet or turbine aircraft includes a electrical pressure sensor of the type which does not include interconnected mechanical linkages, such as capacitive transducers, and which are therefore stable over a wide range of temperature and other adverse conditions. The pressure sensor is coupled to detect the pressure resulting from the operation of the aircraft engine. The system detects departures from normal operating conditions, such as a flameout of the aircraft engine, and restores the aircraft engine to normal operating conditions by reignition, by operating the starter motor, or other similar steps. Controllable hysteresis may be introduced into the system to determine both the pressure at which the control action is initiated and a different pressure where it is stopped. The system may operate in conjunction with pressure metering and other aircraft systems without hysteresis.

Abstract: An aircraft engine which may, for example, be a turbine engine, is provided with an engine temperature thermocouple harness for monitoring the temperature of the aircraft engine. A protective isolation assembly allows for the use of two separate aircraft circuits coupled to receive temperature indicating signals from the thermocouple system. The protective assembly includes two resistors sealed in a metal housing to be connected at the branch point from the existing aircraft temperature monitoring system; and the housing for the protective assembly is electrically and mechanically connected in direct proximity to the thermocouple harness output terminal housing. A short circuit of the second engine temperature monitor circuit would have negligible effect on the existing temperature indicator circuitry in view of the presence of the isolation block assembly.

Abstract: A diesel engine which may, for example, be a V-12 engine, is provided with a turbocharger having two sections, a turbine section which receives exhaust gases and which drives the associated compressor section. The turbocharger increases the input air pressure to the input manifold of the diesel engine up to several times atmospheric pressure. Both the turbine section and the compressor section of the turbocharger have adjustable vanes to vary the coupling with the exhaust gases, in the case of the turbine section, and with respect to the input air in the case of the compressor. An associated control circuit senses the temperature at the inlet to the turbine (the engine exhaust temperature), the speed of the engine, and the engine inlet manifold pressure, and adjusts the orientation of the turbine and compressor vanes for optimal engine performance for all engine speeds, while avoiding cavitation or "stall" conditions in the compressor.

Abstract: A diesel engine which may, for example, be a V-12 engine, is provided with a turbocharger having two sections, a turbine section which receives exhaust gases and which drives the associated compressor section. The turbocharger increases the input air pressure to the input manifold of the diesel engine up to several times atmospheric pressure. Both the turbine section and the compressor section of the turbocharger have adjustable vanes to vary the coupling with the exhaust gases, in the case of the turbine section, and with respect to the input air in the case of the compressor. An associated control circuit senses the temperature at the inlet to the turbine (the engine exhaust temperature), the speed of the engine, and the engine inlet manifold pressure, and adjusts the orientation of the turbine and compressor vanes for optimal engine performance, while avoiding cavitation or "stall" conditions in the compressor.

Abstract: An aircraft having two turbine engines is provided with both torque and temperature sensitive electronic circuitry to control the fuel supplied to the turbine engines. Pushbutton switches are provided for the selection of the desired torque level by the pilot. The fuel to either engine is reduced, regardless of output torque and its relationship to the pushbutton values of selected torque, when the engine temperature exceeds a predetermined safety level. Failsafe operation is provided through the use of a main fuel supply path including a manual fuel control valve extending from the fuel pump to the engine fuel nozzles, and a normally closed bypass fuel control line which selectively diverts fuel from the principal path under the control of the temperature and torque responsive electronic circuitry.

Abstract: Two turbine engines are connected to a common gear box to drive a common load, such as a helicopter rotor, and circuitry is provided for controlling the two mechanically intercoupled engines in order to provide: (1) individual temperature shutoff for the engines, (2) limitation on the maximum total torque supplied by the two engines to avoid exceeding the power handling capability of the gear box for the associated load, (3) speed control for the engines, and (4) substantially equal sharing of the torque load by the two engines. By suitably combining the signals which require analysis, and supplying a single properly conditioned signal to the fuel control unit of each engine, tight control without oscillation or cycling may be achieved. Notch filtering to avoid reinforcement of mechanical resonances, and a variable response characteristic depending on the magnitude of the error signal may be included in the circuitry.

Abstract: The present test apparatus is connected to a jet engine to perform a large number of tests of the mode of operation of the jet engine on both a static and dynamic basis. The multifunctional tester includes four sets of numerical displays on the face of the instrument. The first of these displays shows a sequence number and represents the specific test sequence which is being undertaken by the test apparatus at a particular time. An advance switch is provided to advance to the desired test sequence; and the initiate switch which starts the displayed test sequence. Both the advance switch and the initiate switch have indicator lights associated with them indicating that the test apparatus is ready either to have the sequence advanced, or to initiate a test represented by the sequence number. Input switches are also provided on the face of the test instrument for entering the trim RPM for the particular jet engine under test.

Abstract: Temperature is sensed in equipment where life is reduced by overtemperature. The sensed temperature is compared to a set point and an output indication showing overtemperature operation is provided to the equipment operator. In addition, a signal is produced which is proportional to the sensed temperature above a reference or threshold value. This signal is passed through a coulomb meter which indicates the total charge. This charge is an integration of the product of the time of duration and the amount of overtemperature, for indicating the accumulated thermal stesses.