Abstract: An ion/electrostatic probe for an afterburning gas turbine jet engine comprises an afterburner flameholder (24) electrically insulated from other engine parts by means of non-electrically-conductive support rods (30, 45) or by means of electrically conductive rods (30) insulated (50, 52) from the engine case (12). An electrical connection (54) may be brought through the engine case. The flameholder (24) may be insulated (70) from an attached igniter.

Abstract: An automatic flight control system, for an aircraft having a roll attitude retention outer loop actuator (29), a roll stability inner loop actuator (25) and a control stick (26) for positioning control surfaces of the aircraft to control its roll attitude, includes means (54, 55) to provide a roll error signal (56) indicative of the deviation in roll attitude from a desired roll attitude. The roll attitude retention outer loop actuator (29) is controlled by a proportional (61) and integral (62) function of the roll error when force is not applied to the stick, but only as a proportional function when force is applied to the stick (48, 49, 51, 62). The roll stability inner loop actuator is controlled by a washed out (72) proportional (64) function of the roll error signal to provide short-term roll retention at roll attitudes established by the control stick during turns against trim. Sensing force on the stick (51, 95, FIG.

Abstract: Ions are emitted (14) into the gas stream (15) of a gas turbine engine at a point with respect to the gas stream (15), in comparison with the position of an ion-responsive electrostatic probe (24), so that there is a discernible, reliable difference in the ion impingement on the electrostatic probe (24) during normal gas flow (FIGS. 3 and 4) in contrast with that which occurs during the abnormal gas flow attendant a surge (FIG. 5) thereby to provide early and sustained detection of stall and/or surge conditions in the engine, even in the absence of flame from the combustor (17).

Abstract: A fuel control (23) for an aircraft engine (10) employs super contingency logic (76) in response to low rotor speed of a helicopter (130) engine failure (131) or entry into an avoid region of a flight regime following engine failure (133) to alter (161, 166-169) limits on the gas generator (30) of a free turbine gas engine (10), whereby following engine failure or in periods of extreme power need, risk of stressing an engine to its failure point is undertaken in favor of acquiring enough power to avoid a certain crash.

Abstract: Low reflectivity electrodes (2-4) are formed within recesses extending from a major surface of a semiconductive and piezoelectric substrate (1) (such as gallium arsenide) so as to reduce SAW reflectivity from the electrodes.

Abstract: The blade angle controlling pitch beam servo (26) of a helicopter tail rotor (22) is responsive to a signal manifestation (76) indicative of free turbine engine (20) gas generator speed (78) to provide torque compensation so that the helicopter airframe will not counter-rotate under the main rotor (10) of a helicopter as a consequence of the torque provided thereto by the airframe-mounted engine (20), or in the absence thereof. A trimming embodiment (FIG. 2) provides only sufficient blade angle command (82a) to compensate for that provided by fixed, collective/tail mixing (110-114). Torque compensation tail rotor blade angle commands may be applied through existing stability and autopilot actuators (30-32) or through an additional torque servo (120, FIG. 3).

Abstract: Low cost, easily reproducible piezoelectric deflector (8) operable with low voltages is achieved by providing an interdigital electrode configuration (12, 13) on a surface of piezoelectric material (10) having alternate sites (15, 16) with different piezoelectric response. A preferred embodiment polarizes the sites oppositely in a piezoelectric ceramic substrate (10). Alternate sites in piezoelectric crystals may be removed or have their piezoelectricity substantially reduced. Differential devices (8a, 8b) employ electrodes on both surfaces of the substrate, with a site at one surface being poled oppositely to an adjacent site at the other surface (FIG. 3) or with adjacent sites at both surfaces poled alike, and signal voltages applied oppositely (FIG. 4).

Abstract: A digital fuel control (53) for a helicopter engine (20) controls fuel flow (52) to the engine in response to a turbine reference speed (62) determined in a normal mode (FIG. 5) to be a rated speed, in a fade-in mode (FIG. 6) to be incremented (117, 120) to an estimated optimum minimum speed (114, 115, 125), in an optimizer mode (FIG. 7) to be incremented (138) in a direction (137) leading to least fuel consumption (135), and in a fade-out mode (FIG. 8) to be incremented (151, 153) back to rated speed (154). The invention provides an engine reference speed which results in minimum fuel consumption during cruise flight.

Abstract: In an aircraft automatic flight control system, automatic shutdown (42) of the roll axis thereof (FIG. 1) as a consequence of faults detected therein also provides automatic shutdown of the yaw axis thereof (180) in the same fashion as detection of faults in the yaw axis (171, 167) provides shutdown of the yaw axis. Coupling of the roll axis shutdown into the yaw axis shutdown allows determining which of three axes cause lighting of a single trim shutdown indicator by resetting the yaw shutdown register (145) and simultaneously resetting both the pitch and roll shutdown registers by means of a single cyclic pitch control switch. Automatic yaw trim outer loop and yaw inner loop stability systems are disclosed (FIG. 2).

Abstract: Acoustic Wave Devices comprise lithium niobate substrates 6, 22 having the principal plane cut at zero degrees and 128.degree., respectively, with respect to the Y crystallographic axis thereof, propagating in the X direction, with amorphous silicon dioxide surface layers of 0.42 and 1.24 wavelengths, respectively.

Abstract: In an aircraft automatic flight control system having a reference parameter synchronizing system (70) operable in response to a trim release switch (44), an initial trim release period (139), on the order of a large fraction of a second (137) causes (217, 218) a relatively slow effect trim reference integrator (208, 211) time constant, for smooth transitions of any error signal, followed by a relatively fast (216) effective reference integrator time constant for close, rapid tracking of the reference signal with the actual aircraft parameter.

Abstract: The speed (54, 56) of the free turbine (40) of a helicopter engine (20) is compared (103) with the speed (105, 106) of the helicopter rotor (10) to indicate (101, 102) autorotation, and the deceleration (108) of the rotor above a threshold magnitude (110) is utilized (81, 68, 69) to increase fuel flow (72) to the engine in anticipation of rotor speed droop which would otherwise occur during recovery from the autorotation maneuver.

Abstract: A plurality of thin pressure sensors are made by processing a first large wafer (20, 110) to provide a plurality of electronic devices (28, 122, 124, 125) having a characteristic which varies inversely with strain, and processing a second wafer (40) to provide a plurality of cavities (46) each registered on the second wafer so as to be registerable with a corresponding device on the first wafer. The wafers (20, 40, 110) have thick undoped silicon substrates (21, 41, 114) which are utilized as handles or carriers during the processing, and are stripped off by etching to a highly doped boron etch stop layer (22, 42, 112) when the processing has proceeded to a point where the need therefore has been satisfied. The first wafers (20, 110) are provided with a suitable pattern of borosilicate glass (except in the region where the pressure sensors are formed) so that the two wafers may be joined by a field assisted bonding at a suitable temperature in a vacuum.

Abstract: An electrically controlled (16) mechanical actuator (13) is driven by optically generated (23) electrical power through an optically controlled (28) switch (26) in an electrically isolated module (20). Optical transducers (54, 62, 70) and feedback (29), as well as optical inputs (40, 89, etc.) to electrical signal processing equipment (46, 47, 49) in an electrically isolated module (22) dictate the electrical control (28). A helicopter flight control system is wholly digital, with solar remote power and optical intercommunication.

Abstract: Damping of the primary bending mode of a tower (12) mounting a wind turbine having a control (36) for providing a pitch blade angle reference signal (40) to modulate the pitch of the turbine blades (1) through a pitch change mechanism (38) for constant power is provided by generating the pitch blade angle reference signal as the integral (104) of the summation (266) of a torque/power controlling blade pitch angle reference rate signal (98) with an estimated acceleration signal (255) generated by filtering (250, 252, 254) the blade pitch angle reference signal (40) with the following transfer function ##EQU1##

Abstract: A plurality of thin pressure sensors are made by processing a first large wafer (20, 110) to provide a plurality of electronic devices (28, 122, 124, 125) having a characteristic which varies inversely with strain, and processing a second wafer (40) to provide a plurality of cavities (46) each registered on the second wafer so as to be registerable with a corresponding device on the first wafer. The wafers (20, 40, 110) have thick undoped silicon substrates (21, 41, 114) which are utilized as handles or carriers during the processing, and are stripped off by etching to a highly doped boron etch stop layer (22, 42, 112) when the processing has proceeded to a point where the need therefore has been satisfied. The first wafers (20, 110) are provided with a suitable pattern of borosilicate glass (except in the region where the pressure sensors are formed) so that the two wafers may be joined by a field assisted bonding at a suitable temperature in a vacuum.

Abstract: The difference in the speed (54, 56) of a helicopter gas engine (20), free turbine (40) from a reference speed (62, 64) generates (80) a desired acceleration signal (81). The difference (82) in actual turbine acceleration (84, 86) from desired acceleration is integrated (100) to provide an engine fuel command signal (67-73) whenever (88) the speed error signal exceeds (90) a predetermined threshold magnitude.

Abstract: A four axis force stick provides signals indicative of force applied to the stick in an axis corresponding to a control axis of an aircraft, including pitch, roll, yaw and lift/speed. The force-related signals are applied through proportional and integral gain signal paths to operate electrohydraulic servos that control the aerodynamic surfaces of the aircraft, such as the cyclic and collective blade pitch of the main rotor and the tail rotor blade pitch of a helicopter, or the ailerons, rudder, elevator and thrust of a fixed wing aircraft. Signal conditioning provides a dead band to avoid integrating minute, inadvertent force stick signal outputs, and vernier sensitivity at low forces with high gain at high forces. Analog and digital embodiments are discussed. The relationship between this wholly new mode of aircraft control and ancillary aircraft functions, such as ground steering, autopilot and stability functions, are also discussed.

Abstract: An accelerometer (1) disposed on the support tower (12) of a wind turbine electric generating system in the vicinity of the rotor (10, 16) thereof provides a signal (1) indicative of acceleration of the tower in the direction of the rotor rotational axis. The signal (2) is passed through a band-pass filter (4) for summation (9) with a torque/power controlled (100) blade pitch angle reference rate signal (98), the integral (104) of which provides a blade pitch angle reference signal (40) to control the pitch angle of the rotor blades (10) through a pitch change mechanism (38), thereby to provide additional, positive aerodynamic damping to the tower while modulating blade angle for constant torque/power in response to wind turbulence.

Abstract: In an automatic flight control system, failure of dual inner loop actuators (12, 13) to track position (22, 23) within a permissible tolerance (182) of each other will cause automatic shut down (185, 190) of an outer loop actuator (37) which is responsive (61) to the same proportional signals (54, 55) as are the inner loop actuators. The invention is disclosed in a system in which automatic flight control positioning of the control surfaces of the aircraft is through a fast, limited authority inner loop, the authority of which is centered by repositioning of the outer loop.