Abstract: A protective ring (4) for the fan blades (2, 3) of a gas turbine engine includes a material combination of metal bands and polymer-bonded fiber-weave layers interconnected in alternating arrangement. The protective ring thus formed, which is characterized by tough-elastic properties and low weight, can be used as a full containment (4?) with sufficiently large wall thickness or as a penetration containment (4?) with smaller wall thickness and with a trapping layer arranged on the outer circumference. As a full containment, the protective ring has impact-absorption capacity, as a penetration containment it prevents a further destruction of the broken-through protective ring.

Abstract: This disclosure relates to a method and an arrangement for preventing explosives such as rocket engines from exploding while subjected to an external fire in an ammunition bunker. The basic principle underlying the invention is that the casing surrounding the explosive includes a number of parts which are held together by mounting components made from a shape memory alloy, the properties of which are selected in such a manner that these mounting components no longer hold the various parts of the casing together if the surrounding temperature rises to approach the ignition temperature of the explosive in the rocket engine.

Abstract: A frangible coupling for interconnecting parts comprising a first ring and a second ring coaxially arranged relative to each other and axially joined via an annular array of fuse ligaments equidistantly spaced apart from each other. The ligaments are configured to fail when an abnormal radial load of a predetermined value causes the first and second ring to move out of their coaxial relationship. When all of the fuse ligaments are severed, the communication between the first and second rings is severed. This allows the first ring to move independently of the second ring, preventing the out of balance load on the first ring being communicated to the second ring.

Abstract: An embodiment of the rocket motor of this invention employs an insensitive munitions approach that, when subjected to elevated external temperatures, is activated by thermal expansion of the main propellant and gas generation from a secondary insensitive munitions charge. In a preferred embodiment, the rocket motor also includes a pressure equalizing system that accommodates changing temperature conditions during storage as well as varying gas pressure inherent in gun-launched systems in a manner that allows for thinner case cylinder design and increased propellant volume.

Abstract: A fan thrust reverser includes an outer louver door joined to an inner blocker door by a drive link in a bifold configuration. The louver door is stowed closed in the outer skin of a fan nacelle outside the blocker door stowed closed in the inner skin of the nacelle. An actuator deploys open the louver and blocker doors, with the louver door extending radially outwardly and the blocker door extending radially inwardly for reversing fan exhaust flow during thrust reverse operation.

Abstract: A control strategy for a gas turbine engine which exchanges future lifetime of the engine for present thrust. Gas turbine engines, such as those used in aircraft, sometimes incur damage, as when they ingest birds, or are struck with ballistic objects fired by an enemy. The invention detects the damage, and invokes a control strategy wherein the engine is operated in a more harsh manner, thereby sacrificing a significant part of the remaining lifetime of the engine, in order to obtain thrust currently.

Abstract: Disclosed herein is a “real time” Time Limited Dispatch (TLD) fault management system and method for evaluating the operational suitability of an engine's electronic control system. The TLD system disclosed herein uses a software algorithm to compute or predict the probability of mission success for a given upcoming mission length (e.g., two hours) and the time remaining to repair control system faults before dispatch is disallowed or the flight suspended, regardless of the probability of success.

Abstract: A joint assembly for limiting an extension of the joint in the direction of a load path derived from an impact comprising a first member having a portion and a second member having a portion, the portions overlapping one another and arranged generally parallel to one another and secured together via securing means disposed through corresponding holes defined therein. One of the overlapping portions further defines, sequentially in the direction of extension, a shear neck, a pocket and a catcher portion. In the event of a worst-case impact load the securing means shears through the shear neck and the pocket and is arrested by the catcher portion, thereby the extension of the joint assembly is limited and the joint assembly remains integral.

Abstract: The present invention relates to a method and an arrangement for, in connection with a possible external fire in an ammunition bunker, preventing products container therein comprising encased explosives, such as for example rocket engines, exploding. The basic principle underlying the invention is that the casing (2, 6) surrounding the explosive is to consist of a number of parts which are held together by mounting components (11) made from shape memory alloy, the properties of which are selected in such a manner that these mounting components no longer hold the various parts of the casing together if the surrounding temperature rises to such a great extend that the ignition temperature of the explosive begins to approach.

Abstract: The invention provides a locking system on a cascade thrust reverser for a bypass turbojet, the reverser (20) comprising guide means (60) for the movable parts (50), and the guide means (60) comprising rails (64) attached to the movable parts (50) and tracks (62) attached to the longitudinal members (36).

Abstract: A lock assembly for a thrust reverser system that prevents thrust reverser deployment when the thrust reversers are in the stowed position, but does not prevent thrust reverser movement, in either the deploy or stow directions, when the thrust reversers are out of the stowed position and the lock assembly is in the locked position.

Abstract: The invention provides a device for supporting and recentering a shaft (2) after decoupling. The device comprises an annular support (4) surrounding the shaft (2), a bearing (5) secured to the annular support (4) and having an inner ring (9) which is spaced apart from the shaft by clearance J. The inner ring (9) has two webs or tenons (14a, 14b) that are normally held to the periphery of two troughs (16a, 16b) carried by the shaft (2). Two sets of balls (28a, 28b) normally bear against the bottoms of indentations (25, 26) formed facing one another in the adjacent radial walls of the troughs (16a, 16b) and of the inner ring (9). Resilient means (26) tend to urge the troughs (16a and 16b) towards each other. In the event of the webs or tenons (14a, 14b) breaking due to a large unbalance, the balls (28a, 28b) roll on the walls of the indentations (25, 26) and move the troughs (16a, 16b) apart from each other.

Abstract: An aircraft thrust reverser system that includes a power drive unit operable to supply a drive force. The power drive unit includes a motor and a torque decoupler that sequentially decouples the power drive unit from a plurality of loads. The torque decoupler includes at least two output sections, each of which is coupled to the motor, and are not coupled to one another under normal operating circumstances. The output sections are each operable to decouple the motor from associated thrust reverser moveable components upon a torque magnitude being reached in the output section. The torque decoupler is configured so that the first and second output sections are coupled together a deadband time period after the torque magnitude is reached in on of the output sections.

Abstract: A bearing assembly for a gas turbine engine rotor includes a paired race, a bearing, a mounting joint, and a plurality of mechanical fuses. The paired race includes an outer race and an inner race. The bearing is between the inner and outer races, and is configured to support the rotor on a support frame. The mounting joint includes a joint inner race and a joint outer race. At least one of the joint inner race and the joint outer race includes a spherical surface. The mounting joint is configured to reduce dynamic loads to the gas turbine engine structure and static bending to the rotor. The plurality of mechanical fuses extend through at least one of the joint inner race and the joint outer race.

Abstract: A frangible coupling for interconnecting parts comprising a first ring and a second ring coaxially arranged relative to each other and axially joined via an annular array of fuse ligaments equidistantly spaced apart from each other. The ligaments are configured to fail when an abnormal radial load of a predetermined value causes the first and second ring to move out of their coaxial relationship. When all of the fuse ligaments are severed, the communication between the first and second rings is severed. This allows the first ring to move independently of the second ring, preventing the out of balance load on the first ring being communicated to the second ring.

Abstract: A fan containment case assembly 30 of a gas turbine engine 10 includes a hardened fan case liner 42 disposed therein. In the event of a fan blade loss condition, the hardened fan case liner allows for circumferential movement of the fan blade tips 38 around the fan case 48. Thus, the liner reduces the destructive cutting away of the fan case, minimizing damage to the fan case and decreasing torque loading of the fan case from rotor deflections. Alternate embodiments of the fan case liner are described.

Abstract: A control strategy for a gas turbine engine which exchanges future lifetime of the engine for present thrust. Gas turbine engines, such as those used in aircraft, sometimes incur damage, as when they ingest birds, or are struck with ballistic objects fired by an enemy. The invention detects the damage, and invokes a control strategy wherein the engine is operated in a more harsh manner, thereby sacrificing a significant part of the remaining lifetime of the engine, in order to obtain thrust currently.

Abstract: An embodiment of the rocket motor of this invention employs an insensitive munitions approach that, when subjected to elevated external temperatures, is activated by thermal expansion of the main propellant and gas generation from a secondary insensitive munitions charge. In a preferred embodiment, the rocket motor also includes a pressure equalizing system that accommodates changing temperature conditions during storage as well as varying gas pressure inherent in gun-launched systems in a manner that allows for thinner case cylinder design and increased propellant volume.

Abstract: An embodiment of the rocket motor of this invention employs an insensitive munitions approach that, when subjected to elevated external temperatures, is activated by thermal expansion of the main propellant and gas generation from a secondary insensitive munitions charge. In a preferred embodiment, the rocket motor also includes a pressure equalizing system that accommodates changing temperature conditions during storage as well as varying gas pressure inherent in gun-launched systems in a manner that allows for thinner case cylinder design and increased propellant volume.

Abstract: The present invention is directed to methods and apparatus of detecting engine ignition. The invention generally includes a camera assembly to detect electromagnetic spectra, and a control assembly operatively interconnected with the camera assembly to monitor and evaluate the spectral data collected by the camera assembly. The control assembly is generally programmed with first and second predetermined thresholds associated with respective first and second electromagnetic spectra, and a time threshold at least generally associated with a first time parameter in which at least one of the first and second predetermined thresholds should be reached.

Abstract: A control strategy for a gas turbine engine which exchanges future lifetime of the engine for present thrust. Gas turbine engines, such as those used in aircraft, sometimes incur damage, as when they ingest birds, or are struck with ballistic objects fired by an enemy. The invention detects the damage, and invokes a control strategy wherein the engine is operated in a more harsh manner, thereby sacrificing a significant part of the remaining lifetime of the engine, in order to obtain thrust currently.

Abstract: A gas turbine engine fan blade containment assembly (38) comprising a generally cylindrical, or frustoconical, metal casing (40) has an upstream portion (56), a blade containment portion (54) and a transition portion (58). The upstream portion (56) has a flange (52) connecting the metal containment casing (40) to an axially adjacent casing. The upstream portion (56) of the casing (40) has a greater diameter (D1) than the diameter (D2) of the blade containment portion (54) and the blade containment portion (54) has a greater thickness (T2) than the thickness (T1) of the upstream portion (54). The transition portion (58) connects the blade containment portion (54) and the upstream portion (56) to transmit loads from the blade containment portion (54) to the upstream flange (42).

Abstract: A ducted fan gas turbine engine comprises a fan 12 mounted on a first shaft 16 and a compressor mounted on a second shaft 43. The shafts are coaxial both with each other and with said engine longitudinal axis 15. The shafts are mounted in roller bearings 21, 40. The first bearing 21 carrying the upstream part of the first shaft and the second bearing 40 carrying the upstream part of the second shaft. Both bearings are attached to a structure 26 which comprises an axial sleeve and a front frustoconical portion and a rear radially stiff but axially flexible frustoconical portion and is attached to the engine casing 36. The front frustoconical portion is frangibly attached to the front part of the axial sleeve 28. In the event of damage to the fan the frangible connection breaks and the first shaft is allowed to orbit about the engine longitudinal axis 15. The rear member 32 is swashingly non-linear flexible and accommodates the out of balance radial loads transferred to it as couples during such an incident.

Abstract: A microthruster having an inverted exhaust system traps burst diaphragm fragments providing a clean exhaust while an exhaust port provides increased back pressure for efficient combustion of a propellant charge in a fuel cell. A converging diverging micronozzle provides a predictable exhaust vector for improved microthrusting well suited for propulsion system on small spacecraft.

Abstract: A bearing support includes a hollow cone diverging between a forward bearing seat and an aft mounting flange. The cone includes an integral annular fuse sized in thickness to fail in shear under abnormal imbalance load transfer between a gas turbine engine fan supported by a bearing in the forward seat and a fan frame mounted to the aft flange.

Abstract: A rotor assembly for a gas turbine engine including a bearing assembly and a damper sub-assembly that facilitate reducing dynamic motion to the rotor assembly is described. The bearing assembly includes rolling elements positioned between a paired race. The rotor assembly includes a rotor shaft supported by the bearing assembly. The damper sub-assembly is radially outward from the bearing assembly adjacent a sump housing, and includes a damper insert. A predetermined preload force is applied to the rolling elements. An outer race of the bearing assembly distorts to substantially match a distortion pattern of the damper insert.

Abstract: An ordnance venting system has a thermally initiated rapid deflagrating composition that detonates a high energy material. The high energy material is configured as a shaped charge and surrounds the ordnance housing of a rocket motor. During cook-off of the munition, the high energy material detonates, rupturing the structural integrity of the ordnance housing sufficiently to release pressure therefrom.

Abstract: A rocket motor having a desensitizing mechanism for preventing explosion or violent reaction during slow cook-off is provided. The rocket motor includes a case contained rocket propellant with a desensitizing assembly attached forward of the propellant charge. The desensitizing assembly is formed with an enclosure containing a desensitizing fluid, and connected to the interior of the rocket motor by a tube which is sealed by a heat sensitive plug. The heat sensitive plug melts at a temperature below the slow cook-off temperature of the rocket propellant. The melting of the plug allows the desensitizing fluid to be ejected into the interior of the casing and onto the propellant charges, thereby, desensitizing the propellant charge.

Abstract: A mount for a gearbox accessory for a gas turbine engine substantially isolates the accessory from direct mechanical loading from the engine case during normal engine operating conditions and couples the accessory to the engine case during high rotor imbalance conditions to minimize interface loads between the gearbox and the accessory. By coupling the gearbox to the accessory, the two are displaced substantially in unison, thereby minimizing interface stresses therebetween. Various construction details of the mount for the gearbox accessories are disclosed.

Abstract: In a control system for a gas turbine aero engine, the control system (ECU) is configured as a dual control system comprising two channels, Ch-A and Ch-B. Ch-A has two CPUs which conduct calculations separately based on the sensor outputs and one of the CPUs compares the results and if they coincide, the CPU sends the result of the other CPU to the FCU. If not, one of the CPUs determines that an abnormality arises in Ch-A and sends a result to Ch-B. Ch-B is constituted as a standby channel having only one CPU whose operation is monitored by a simple watchdog timer circuit. This enhances CPU failure detection with a relatively simple configuration, and eliminates the need for provision of an overspeed protector.

Abstract: To miniaturize an exhaust manifold to facilitate disposition of an engine in a boat body. A multiple cylinder engine having an exhaust manifold on a side portion thereof is disposed such that a crankshaft thereof extends along forward and rearward directions of a boat body. The exhaust manifold has a structure wherein exhaust pipes connected to exhaust ports of the engine are distributed upwardly and downwardly in such a manner as to form confluence pipes above and below the exhaust ports and the confluence pipes are further joined to each other to form an exhaust gas exit from the exhaust manifold. The exhaust gas exit is disposed at a substantially central portion of an internal space of the boat body in the vertical direction and an exhaust gas turbo charger is connected to the exhaust gas exit.

Abstract: A deployment inhibiting arrangement for a jet engine thrust reverser system, the system including one or more actuators for operating a thrust reverser of the system and first locking means for the or each actuator normally serving to hold the system in a condition in which the thrust reverser is undeployed, the inhibiting arrangement including additional locking means normally held in an inactive condition, but operable by means independent of said one or more actuators and said first locking means to provide an additional mechanical lock serving to prevent deployment of the thrust reverser irrespective of the condition of the first locking means and/or actuator(s). The invention also relates to a jet engine thrust reverser system, the system including such a deployment inhibiting arrangement.

Abstract: An ordnance venting system has an ordnance device having a casing with a vent opening, a dome plug fitted into the formed vent opening and an adapter fitted over the dome plug on the outside of the casing. The adapter connects sufficiently to the casing to retain the dome plug against the formed vent opening for given pressures. The adapter melts at high temperatures and releases the dome plug to reduce the danger of explosion from heat induced over-pressurization.

Abstract: A rocket motor having a desensitizing mechanism for preventing explosion or violent reaction during slow cook-off is provided. The rocket motor includes a case contained rocket propellant with a desensitizing assembly attached forward of the propellant charge. The desensitizing assembly is formed with an enclosure containing a desensitizing fluid, and connected to the interior of the rocket motor by a tube which is sealed by a heat sensitive plug. The heat sensitive plug melts at a temperature below the slow cook-off temperature of the rocket propellant. The melting of the plug allows the desensitizing fluid to be ejected into the interior of the casing and onto the propellant charges, thereby, desensitizing the propellant charge.

Abstract: Plugs for protecting internal spaces of rocket combustion chambers against moisture, dust, and other environmental factors. Each plug includes a central rod with a bracket fastened to it. The plug also includes a sleeve with radially extending structural ribs. Spring-loaded cams with rollers are securely hinged at the ends of the ribs. In addition, a spring-loaded limiter and a valve are mounted on the central rod so that they can possibly move relative to the central rod. The limiter includes a sleeve and a conical shell that are connected by rod ties. The valve includes a metal stamped conical plate that forms a gas-tight fit with the rod. A sealing ring of elastic material is positioned on a peripheral part of the plate. The set of spring-loaded cams and the limiter secure the self-mounting plug in a working position and make the valve operation independent of the operation of other plug components.

Abstract: A support structure for a gas turbine engine includes a support arm (214, 216) extending between a low pressure shaft (202) and a rotor. The support arm includes a fuse (220) having a low failure point. A high pressure stub shaft (270) axially and radially supports the low pressure shaft (202) after fuse failure. An axial gap (A) between a portion of the low pressure shaft and the stub shaft permits movement of the low pressure shaft after fuse failure. A radial gap (B) between the stub shaft (270) and the low pressure shaft (202) allows radial deflection of the low pressure rotor system after fuse failure.

Abstract: A system for controlling aeromechanical instability or flutter in turbofan engines having fan blades employs a sensor, such as an off-blade static pressure sensor or proximity detector mounted on a turbofan engine at an inlet of a rotor of the engine for generating a signal to detect resonance of the turbofan blades at frequencies associated with flutter. A controller is coupled to the sensor for generating by spatial Fourier decomposition from the sensor signal a command signal comprising a real time amplitude component and a spatial phase of disturbances of a predetermined nodal diameter and coincident with a natural frequency of resonance of a predetermined structural mode of the fan blades in the stationary frame. An actuator, such as a bleed valve or acoustic speaker, is mounted on the turbofan engine for damping flutter dynamics in response to the amplitude of the command signal.

Abstract: A locator for a gearbox mount for a gas turbine engine includes a mounting plate having a first stub shaft that extends from the plate into a receptacle in the gearbox, a second stub shaft that extends from the plate into a receptacle in the engine case and a load sensitive mechanism disposed in either one of the stub shafts. During a high rotor imbalance event such as a fan blade loss condition, the load sensitive mechanism shears to partially decouple the gearbox from the engine case to minimize damage caused by high rotor imbalance loads. Alternate embodiments of the load sensitive mechanisms in the stub shaft are disclosed, including a diaphragm and a wall that yield to shear loads.

Abstract: The present invention relates to a low-vulnerability solid-propellant motor (1) comprising a charge (11) of solid propellant, a heat shield (2) covering the inside of a metal structure with a front end wall (3) and a rear end wall (4) which carries at least one nozzle (10), the said front and rear end walls being fastened to a cylindrical shell (5) having wide apertures which are circumferentially distributed and temporarily sealed by sealing plugs (8). The object of the present invention is to provide a motor which has a simple structure and avoids the use of detection and actuation means for freeing the plugs. The motor according to the invention is such that the sealing plugs (8) are held in the apertures between the heat shield and a composite winding (7) of fibres impregnated in a matrix which surrounds the cylindrical shell (5).

Abstract: A lock for a thrust reverser mechanism having actuator means for operating a thrust reverser mechanism, includes a first lock device connectable, in use, to the actuator means for preventing actuation of the latter. The first lock device is operable hydraulically to unlock the lock so as to free the actuator means for operation. An electro-mechanical second lock device is arranged so as normally to prevent actuation of the first lock device and is operable to permit actuation of the first lock device when required.

Abstract: A ducted fan gas turbine engine (10) is provided with a fan (12) which is carried by a shaft (16). A frangible support member (27) maintains a coaxial relationship between the shaft (16) and the longitudinal axis (15) of the engine (10). In the event of severe damage to the fan (12), the frangible member (27) fractures, allowing the shaft (16) carrying the fan (12) to orbit about the engine longitudinal axis (15). A resilient support member (33,45) exerts a radially inward restoration force on the shaft (16) carrying the fan (12) to reduce vibration.

Abstract: A ducted fan gas turbine has a fan shaft which is supported at its upstream end by a bearing. In the event of damage to the engine's fan, frangible bolts fracture to permit radial movement of the shaft. That movement is limited by the sequential engagement of the outer race of the bearing with first and second movement limiting portions on fixed structure of the engine. This alters the natural frequency of vibration of the fan, so permitting the fan to windmill without excessive vibration.

Abstract: One embodiment of this novel rocket system employs at least one rocket motor closure that includes low and high load capability retainers that, respectively, provide an insensitive munitions system during storage and positive locking of a nozzle assembly at lunch. Another embodiment of this novel rocket involves an insensitive munitions approach that does not require use of such retainers by developing different normal and abnormal burn paths. The rocket motor also includes a pressure equalizing system that accommodates changing temperature conditions during storage as well varying gas pressure inherent in gun-launched systems in a manner that allows for thinner case cylinder design and increased propellant volume.

Abstract: A ducted fan gas turbine engine is provided with a fan which is carried by a shaft. Frangible fuse pins maintain the shaft carrying the fan coaxial with the engine longitudinal axis. In the event of severe fan damage, the fuse pins fracture and the fan shaft orbits about the engine longitudinal axis. An annular resilient member is provided to exert a radially inward restoration force upon the fan shaft following such an event in order to reduce vibration.

Abstract: A solid fuel propellant motor casing is described, the casing comprising a body portion for containing the propellant, the body portion also including at least one detachable end closure portion, the closure portion being maintained in position relative to said body portion by retaining means, the retaining means being held in position with engagement means on an inner surface of said body portion and located radially with respect to said body portion by locating means against resilient biasing means, said locating means being responsive to temperature such that said locating means melt at a predetermined temperature allowing said resiliently biased retaining means to move radially inwardly of said body portion and allow said at least one closure portion to be ejected from the casing.

Abstract: A ducted fan gas turbine engine is provided with a fan which is carried by a shaft. A frangible fuse ring maintains the shaft carrying the fan coaxial with the engine longitudinal axis. In the event of severe fan damage, the fuse ring (28) fractures and the fan shaft orbits about the engine longitudinal axis. A pair of threaded members (39, 41), one of which is of generally frusto-conical configuration, cooperate as a result of the orbiting motion to restore the coaxial relationship between the fan shaft and the engine longitudinal axis.

Abstract: A method is described for enabling operation of an aircraft turbo-engine to be continued in spite of the accidental occurrence of an unbalance in a tor of the engine, such as may be caused by the breakage of a fan blade. The method involves constructing the bearing support of at least one of the bearings of the rotor so that the support has at least two levels of stiffness, arranging for the support to change from a normal stiffness level to a lower stiffness level in response to the occurrence of an unbalance in excess of a pre-determined minimum value, and setting the clearance between the tips of the rotor blades and the inner wall of the surrounding casing so that it is at least equal to the radial displacement of the rotor under the action of a pre-determined maximum unbalance when the rotor is rotating at the flight idling speed of the engine and the bearing support is at the lower stiffness level. A bearing support construction for implementing the method is also described.

Abstract: A device for initiating burning of an explosive or propellant material subjected to external heating prior to detonation is described, the device comprising at least two materials, at least a first one of which materials melts at a temperature below the temperature at which detonation of said explosive or propellant material is induced and reacts with said second material to produce an exothermic reaction to cause initiation of burning in said explosive or propellant material.

Abstract: A responder mounted to the exhaust nozzle support structure of a gas turbine engine includes a pressure switch and a sensor tube connected thereto. The sensor tube encircles the nozzle flaps seals in spaced relation thereto, and the pressure switch is connected to a fire warning alarm.

Abstract: A fencing assembly (66) for prohibiting circulation of primary and/or fan airflows (52), (54) into a bounded low pressure region of a jet engine installation. The fencing assembly includes one or more flow control fences. In an exemplary application, the fencing assembly includes a number of fences positioned around the sides of a batcave (48) bounded low pressure region that is located between a primary exhaust nozzle (20) and strut fairings (30). The fencing assembly includes five fences, each fence being connected to either the primary exhaust nozzle or the strut fairings. The five fences include single fences (70), (72) positioned along each lateral side of the batcave; two rear fences (74), (74') positioned circumferentially about rear regions of the batcave; and an arcuate fence (76) positioned near the two rear fences to form a half circle. The fences include a foot portion (78) and an upright portion (80).