Abstract: A comprehensive turbine engine monitoring and recording system includes an electronics unit which may be engine mounted, and which contains at least a nonvolatile memory and a data-processor, for use with a remote unit suitable for either cockpit mounting or for accessing by a ground portable unit which may include an additional data-processor and display and/or printer units. Raw input data is supplied to the electronics unit, and this input information may include engine and outside air temperature, torque, engine speed, vibration, altitude and airspeed data, for examples. Exceedance data relative to the operation of the aircraft above pre-defined limits of temperature, torque and speed are recorded, with the details of each exceedance being permanently stored.

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 thermocouple system has been nicknamed the "Binocular" thermocouple assembly because the terminal housing has a configuration similar to a pair of binoculars, with the chromel and alumel terminals mounted in spaced circular openings in the terminal housing. A thermocouple probe extends outwardly from the terminal housing, and is provided with two peripheral grooves, into one of which the terminal housing is secured by swaging. The second groove in the thermocouple probe is spaced away from the terminal housing by a short distance along the probe, and a washer is securely swaged into this second groove. A nut having external threads is mounted between the terminal housing and the washer for securing the thermocouple probe to its sensing location, for example, mounted adjacent a turbine engine.

Abstract: A helicopter turbine engine over-stress warning and protection system includes arrangements for sensing the engine temperature, the engine speed, and the output torque from the engine. An audible signal warning is provided which may vary as the over-stress limits are approached. For example, when the helicopter engine is at or above the over-stress limits, it may have a constant fairly loud tone into the pilot's earphones or into a speaker in the cockpit. This warning tone signal becomes interrupted and provides progressively greater periods of silence as the helicopter engine drops below the over-stress limits. The warning signal stops at a pre-set point such as, for example, when the engine is being operated from 5% to 15% below the over-stress limits.

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: A turbine engine monitoring and recording system includes an engine mounted unit which contains at least a non-volatile memory and a data-processor, and a remote unit suitable for either cockpit mounting or for accessing by a ground portable unit including an additional data-processor and a display unit. Thermocouples and other sensors mounted on the engine supply raw data to the engine mounted electronics unit, and this input information includes temperature data. Elevated over-temperature levels are segregated and the time during which the engine is within each of the over-temperature bands is measured, and is recorded in a non-volatile, electronically alterable memory which is periodically updated during operation of the aircraft. The number of starts is also counted and stored, and the total running time of the engine is also recorded.

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 turbine engine may occasionally be required to operate momentarily at an overtemperature condition which, provided it does not exceed a predetermined maximum temperature and time duration, is known to be harmless. A temperature read-out device attached to an engine operating under such conditions would measure this harmless overtemperature transient, indicate an overtemperature state, and display a warning. The circuitry of the present invention monitors a signal representing engine temperature, and beginning at the start of the harmless overtemperature transient, and generates an opposing compensating signal for a predetermined period of time. The opposing compensating signal is combined with the signal representing engine temperature and the resultant signal, is applied to the temperature readout device so that the signal representing engine temperature is reduced below the level which would cause an overtemperature condition to be indicated on the temperature readout device.

Abstract: Pushbutton controls are provided for the power management of a turbine powered aircraft; and these pushbuttons may be mounted on the aircraft pilot's control handwheel. The turbine engine has a maximum rated permissible rotational speed which initially increases with increasing air temperature and with increasing altitude or reduced pressure; and has an absolute maximum limitation, with this maximum permissible rotational speed decreasing at increasing temperatures starting at about 10 or 15 degrees below zero, centigrade; and these limitations are reduced when supplemental equipment such as de-icing equipment is turned on. In accordance with the present invention, a series of "maps", or rotational speed control characteristics reflecting the factors mentioned above, are provided, and the pushbutton controls select among these characteristics, with the "take-off" power button permitting the highest maximum speeds, etc.

Abstract: The maximum speed of a turbofan engine varies in accordance with moderately complex relationships involving the temperature, the pressure or altitude, and whether or not the subordinate systems such as the anti-ice systems are operative. The maximum allowable rotational speed for the fan section of a turbofan engine increases with increased altitude or reduced pressure, and normally increases as the ambient temperature increases up to speeds limited by the centrifugal and other forces acting on the turbine blades. However, there is a maximum fan rotational speed above which the engine should not be operated, and there is a limit which decreases with increasing temperature above which the fan should not be operated regardless of the pressure. In addition, when the anti-ice or other subordinate systems are operative, these limits are reduced substantially.

Abstract: A high temperature thermocouple wiring harness and connector system incorporating multiple individual thermocouple probes replaceably attached through connectors to the wiring harness. Each connector has a wiring harness portion, a removable probe portion, and a fastener for holding the two portions together. The wiring harness portion has a housing and two electrical conductors connected to the cable of the wiring harness and having contact surfaces outside the housing. The removable probe portion has a similar housing, a thermocouple probe mounted on the housing, and two electrical conductors attached to the thermocouple and having contact surfaces outside the housing. The fastener retains the connector portions together and insures that the contact surfaces of the conductors are held together. The fastener is formed of a material that has a coefficient of thermal expansion less than the coefficient of thermal expansion of the connector portions.

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.