Abstract: A gas turbine engine includes a combustor having a combustor air inlet, an axial-centrifugal compressor, a shroud, a secondary flow duct, and a valve. The shroud surrounds at least a portion of the axial-centrifugal compressor and has a surge bleed plenum defined therein that is in fluid communication with, and receives compressed air from, the axial compressor outlet. The secondary airflow duct has a duct inlet that is in fluid communication with the surge bleed plenum, and a duct outlet that is in fluid communication with the combustor air inlet. The valve is mounted on the secondary airflow duct and is movable between a closed position, in which the secondary airflow duct does not provide fluid communication between the surge bleed plenum and the combustor air inlet, and an open position, in which the secondary airflow duct provides fluid communication between the surge bleed plenum and the combustor air inlet.

Abstract: The device comprises two heat exchangers (74, 90) suitable respectively for vaporizing first and second propellants before they are reintroduced in gaseous form into their tanks (16, 18). The heat exchangers co-operate respectively with first and second gas generators (60, 84) suitable for being fed with a mixture of propellants in order to produce combustion, the second gas generator (84) being suitable for being fed at least in part by the exhaust from the first gas generator (60).

Abstract: An assembly is provided for a turbine engine. The assembly includes a cooling tube through which a flow path extends. The assembly also includes a fuel injector, which includes a stem connected to a tip. The stem is arranged within the cooling tube. The tip extends from the cooling tube to a nozzle. The cooling tube is adapted to direct cooling air through the flow path.

Abstract: A combustor for a gas turbine engine comprising a combustion liner defining a combustion chamber, wherein the combustion liner has at least one opening into which a combustor liner mount is received. The combustor liner mount can have at least one cooling passage defined therein having an inlet defining an inlet axis and an outlet defining an outlet axis, wherein the inlet axis is not collinear with the outlet axis for providing enhanced cooling to at least one of the combustor liner mount and an adjacent combustion component, such as an igniter or a borescope plug.

Abstract: A control system for a gas turbine engine, a method for controlling a gas turbine engine, and a gas turbine engine are disclosed. The control system may include a hybrid model predictive control (HMPC) module, the HMPC module receiving power goals and operability limits and determining a multi-variable control command for the gas turbine engine, the multi-variable control command determined using the power goals, the operability limits, actuator goals, sensor signals, and synthesis signals. The control system may further include system sensors for determining the sensor signals and a non-linear engine model for estimating corrected speed signals and synthesis signals using the sensor signals, the synthesis signals including an estimated stall margin remaining. The control system may further include a goal generation module for determining actuator goals for the HMPC module using the corrected speed signals and an actuator for controlling the gas turbine engine based on the multivariable control command.

Abstract: A control system for a gas turbine engine, a method for controlling a gas turbine engine, and a gas turbine engine are disclosed. The control system may include a hybrid model predictive control (HMPC) module, the HMPC module receiving power goals and operability limits and determining a multi-variable control command for the gas turbine engine, the multi-variable control command determined using the power goals, the operability limits, actuator goals, sensor signals, and synthesis signals. The control system may further include system sensors for determining the sensor signals and a non-linear engine model for estimating corrected speed signals and synthesis signals using the sensor signals, the synthesis signals including an estimated stall margin remaining. The control system may further include a goal generation module for determining actuator goals for the HMPC module using the corrected speed signals and an actuator for controlling the gas turbine engine based on the multivariable control command.

Abstract: A combustor nozzle assembly includes: a nozzle mounting base which blocks a combustor insertion opening formed in a turbine casing; a nozzle rod which passes through the nozzle mounting base and has a rod tip portion and a rod base end portion; an oil fuel pipe which is as a whole inserted into the nozzle rod, which has a pipe tip portion and a pipe base end portion, in which fuel is supplied to the inside through the rod base end portion, and which injects the fuel from the pipe tip portion through the rod tip portion; and an O-ring which is disposed in the rod base end portion and suppresses leakage of fuel to the pipe tip portion side between the inner periphery side of the nozzle rod and the outer periphery side of the oil fuel pipe.

Abstract: A control system for a gas turbine engine, a method for controlling a gas turbine engine, and a gas turbine engine are disclosed. The control system may include a hybrid model predictive control (HMPC) module, the HMPC module receiving power goals and operability limits and determining a multi-variable control command for the gas turbine engine, the multi-variable control command determined using the power goals, the operability limits, actuator goals, sensor signals, and synthesis signals. The control system may further include system sensors for determining the sensor signals and a non-linear engine model for estimating corrected speed signals and synthesis signals using the sensor signals, the synthesis signals including an estimated stall margin remaining. The control system may further include a goal generation module for determining actuator goals for the HMPC module using the corrected speed signals and an actuator for controlling the gas turbine engine based on the multivariable control command.

Abstract: A gas turbine engine fuel delivery system which includes an external fuel manifold fastened to the outer surface of a casing surrounding the combustor of the engine and which includes a plurality of inlet manifold tubes circumferentially disposed in serial flow communication. A plurality of fuel nozzles are mounted to the casing and are axially spaced apart from the fuel manifold on the outer surface of the casing. A plurality of jumper tubes are disposed outside the casing and feed fuel flow from the fuel manifold to the fuel nozzles. The jumper tubes have a rigidity that is less than that of the inlet manifold tubes of the fuel manifold such that the jumper tubes elastically deflect under load.

Abstract: According to one aspect of the invention, a gas turbine engine includes a combustor, a fuel nozzle placed in an end of the combustor, and a passage configured to receive an air flow from a compressor discharge casing, wherein the passage directs the air flow into a chamber downstream of the nozzle, wherein a chamber pressure is lower than a compressor discharge casing pressure. The gas turbine engine also includes a flow control device configured to control the air flow from the compressor discharge casing into the passage.

Abstract: A system includes a controller configured to control a semi-closed power cycle system. The controller is configured to receive at least one of a first signal indicative of an oxygen concentration within a first gas flow through a primary compressor, a second signal indicative of power output by the semi-closed power cycle system, a third signal indicative of a temperature of a second gas flow through a turbine, and a fourth signal indicative of a mass flow balance within the semi-closed power cycle system. The controller is also configured to adjust at least one of the first gas flow through the primary compressor, a fuel flow into a combustor, a fraction of the first gas flow extracted from the primary compressor, and an air flow through a feed compressor based on the at least one of the first signal, the second signal, the third signal, and the fourth signal.

Abstract: The device for injecting a liquid mono-propellant with a large amount of modulation of its flow rate and disposed at an upstream end of the wall of a combustion chamber of a rocket engine has a feed channel for feeding a mono-propellant from a tank. The device includes a single annular speed-up channel connected to the feed channel and having its outlet opening out via an annular injection section, the speed-up channel and the annular injection section being defined firstly by a first wall forming a stationary surface of revolution situated level with said upstream end, and secondly by a second wall forming a surface of revolution that is on a part that is movable in translation relative to the first wall forming a stationary surface of revolution.

Abstract: Apparatus for reducing air mass flow through the compressor in a single shaft gas turbine engine having an extended operating range including part load conditions, to provide low emissions combustion. The apparatus includes one or more nozzles positioned for injecting compressed air into the inlet region of the compressor. The nozzles are oriented to direct the compressed air tangentially to, and in the same angular direction as, the direction of rotation to create a swirl in the inlet air flow to the compressor inducer. The apparatus also includes conduits in flow communication between the compressor diffuser and the nozzles, one or more valves operatively connected to control the flow of compressed air from the diffuser to the nozzles, and a controller operatively connected to the valves to cause compressed air flow to the nozzles during operation at part load conditions.

Abstract: A fuel controller, and associated method, provides a fuel control output signal to a fuel control actuator to control operations. The fuel controller determines the fuel control output signal based on rotational speed error. A combustion air controller provides a combustion air control output signal to a combustion air control actuator to control operations. A cross channel controller is in communication with the fuel controller and the combustion air controller. The cross channel controller provides a combustion air control modification signal to the combustion air controller. The combustion air control modification signal is determined from the fuel control output signal using an air versus fuel model. The combustion air controller determines a preliminary combustion air control signal based on an exhaust temperature error, and further determines the combustion air control output signal based on both of the preliminary combustion air control signal and the combustion air control modification signal.

Abstract: A coordinated air-fuel controller and associated method provide a fuel controller, a combustion air controller and a steady-state air versus fuel model. The fuel controller generates a fuel control output signal and the combustion air controller generates a combustion air control output signal. The fuel controller determines a preliminary fuel control signal based on at least one of first and second loop control signals, and determines the fuel control output signal based on the preliminary fuel control signal. The steady-state air versus fuel model processes the preliminary fuel control signal to determine an expected steady-state combustion air control signal. The combustion air controller determines a preliminary combustion air control signal based on at least one of a third loop control signal and a fourth loop control signal, and determines the combustion air control output signal based on the preliminary combustion air control signal and the expected steady-state combustion air control signal.

Abstract: A compressor includes a diffuser, a recirculation duct, and a flow control valve. The recirculation duct has an inlet in fluid communication with the air outlet of the diffuser, and an outlet in fluid communication with the air inlet of the diffuser. The flow control valve is selectively moveable between open and closed positions, to thereby fluidly couple and isolate, respectively, the recirculation duct inlet and outlets. During operation of the compressor, the flow control valve may be opened, which circulates a portion of the compressed air discharged from the diffuser air outlet back to the diffuser air inlet. The air that is circulated back to the diffuser air inlet reduces the effective area of the diffuser air inlet, thereby increasing the surge margin of the compressor.

Abstract: An oxygen-hydrogen pressurization system includes a cryogenic oxygen tank, cryogenic hydrogen tank, thermal switch, supercritical oxygen bottle, supercritical hydrogen bottle, and pressure management system and a thermodynamic vent system. The thermal switch permits heat to flow between hot and cool areas within the space vehicle to help facilitate pressure management within the cryogenic liquid oxygen tank and the cryogenic liquid hydrogen tank in conjunction with the higher pressure fluid from the supercritical oxygen tank and the fluid from the supercritical hydrogen tank and the added cooling from the pressure management system.

Abstract: A liquid/gas interface pressure accumulator and mechanical filter device is especially adapted for use within a propellant utilization monitoring system of a space launch vehicle. The device provides a retrofit solution for overcoming problems associated with liquid entering a pressure sensing lines and pressure noises that collectively harm the ability to accurately measure fuel and oxidizer levels. One device is attached to each fuel and oxidizer tank of the launch vehicle. Each device includes a housing and an internal sensing chamber that is used to accumulate liquid before it can enter a pressure sensing communication line connected to a pressure sensing transducer. The construction of the device also allows filtering of undesirable pressure noises due to vibration, pressure pulsations, and other vibratory events that occur during operation of the space vehicle. The device is also selectively tunable to filter out certain frequencies and frequency ranges/bands.

Abstract: A method for determining a target exhaust temperature for a gas turbine including: determining a target exhaust temperature based on a compressor pressure condition; determining a temperature adjustment to the target exhaust temperature based on at least one parameter of a group of parameters consisting of specific humidity, compressor inlet pressure loss and turbine exhaust back pressure; and adjusting the target exhaust temperature by applying the temperature adjustment.

Abstract: The thrust of a rocket motor can be varied to optimize Nozzle Pressure Ratio (NPR) using a design that allows for adjusting the relative position of a plug and a combustion chamber exit. The plug or the exit may be attached to an adaptive control system for position modification. The relative position of the plug and exit may be adjusted to optimize NPR to account for changing propellant flow and/or changing ambient pressure.

Abstract: Sliding-Action Magneto-Mechanical Injector Throttling Device (SLAMMIT) provides on-demand, yet accurate, throttling of the mass flow of the fuel and/or oxidizer into the combustion chamber of a vortex injector. At least two SLAMMIT sub-assemblies comprise the SLAMMIT Device and each sub-assembly is integrated into a manifold and is driven to slide in a given direction by a drive block that is internal to the sub-assembly. The drive block is, in turn, actuated by an electromagnet that is external to the SLAMMIT sub-assembly. As the SLAMMIT sub-assemblies slide, flappers inside the sub-assemblies achieve the effective opening size of the injection orifices anywhere between fully open and fully closed.

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 throttleable rocket injector assembly is disclosed wherein the flow of oxidizer into each of the oxidizer injectors is simultaneously regulated by a mechanical assembly mounted entirely within the oxidizer chamber.

Abstract: A method and system are provided for propelling an aerodynamic vehicle into space. The aerodynamic vehicle uses a nuclear-based thermal rocket (NTR) propulsion system capable of producing a hydrogen exhaust. A flow of air is introduced into the hydrogen exhaust to augment the thrust force at speeds of the vehicle up to approximately Mach 6. When the speed of the vehicle is approximately Mach 6 and the altitude of the vehicle is approximately 40 kilometers, the flow of air is stopped and the vehicle is propelled into space using only the NTR.

Abstract: A control system for a ducted fan gas turbine engine (10) comprises a control unit (29) that receives input signals from a pressure transducer (27) located within the engine's fan duct (11) downstream of the fan (13) and a rotational speed transducer (32) mounted adjacent the shaft (24) carrying the fan (13). In the event of the control unit (29) detecting a fall in the air pressure downstream of the fan (13) and an increase in fan speed, it commands a temporary reduction in the fuel flow to the engine (10). Such a decrease in the air pressure downstream of the fan (13) and increase in fan speed is indicative of the fan (13) entering a stall condition. The temporary reduction in fuel flow to the engine (10) enables the fan (13) to recover from that stall condition.

Abstract: A dual concentric poppet valve assembly (44) includes an assembled body (45) having an opening (55) therethrough and having a sealing surface (56) surrounding the opening. A first member (46) is mounted in the body opening for axial movement relative to the body. The first member has a first sealing surface (70) arranged to be moved toward and away from the body sealing surface (56) to define a first variable-area orifice (96) therebetween. The first member has a concentric opening (72) therethrough and has a second sealing surface (71) surrounding the first member opening. A second member (48) is mounted in the first member opening for co-axial movement relative to the first member. The second member has a sealing surface (83) arranged to be moved toward and away from the first member second sealing surface (71) to define a second variable-area orifice (98) therebetween.

Abstract: In a fuel supply system for a space missile or a spacecraft in which the thrusters are supplied with liquid or gaseous fuel from a fuel tank in which the tank pressure decreases continuously during the operational life of the system, the fuel throughput and thus the thrust of the respective thruster or thrusters is maintained substantially constant by the arrangement of a constant flow device in the fuel line between the tank and the reaction chamber of the system. The constant flow device is a self regulating pressure loss device realized by increasing or decreasing the through-flow area of the device in response to changing fuel supply pressures.

Abstract: In the representative embodiment of the new and improved methods and apparatus disclosed herein to provide effective real-time management of a spacecraft rocket engine powered by gaseous propellants, real-time measurements representative of the engine performance are compared with predetermined standards to selectively control the supply of propellants to the engine for optimizing its performance as well as efficiently managing the consumption of propellants. A priority system is provided for achieving effective real-time management of the propulsion system by first regulating the propellants to keep the engine operating at an efficient level and thereafter regulating the consumption ratio of the propellants. A lower priority level is provided to balance the consumption of the propellants so significant quantities of unexpended propellants will not be left over at the end of the scheduled mission of the engine.

Abstract: In the representative embodiment of the new and improved methods and apparatus disclosed herein to provide effective real-time management of a spacecraft rocket engine powered by gaseous propellants, real-time measurements representative of the engine performance are compared with predetermined standards to selectively control the supply of propellants to the engine for optimizing its performance as well as efficiently managing the consumption of propellants. A priority system is provided for achieving effective real-time mangagement of the propulsion system by first regulating the propellants to keep the engine operating at an efficient level and thereafter regulating the consumption ratio of the propellants. A lower priority level is provided to balance the consumption of the propellants so significant quantities of unexpended propellants will not be left over at the end of the scheduled mission of the engine.

Abstract: A rocket injector is provided with multiple sets of manifolds for supplying propellants to injector elements. Sensors transmit the temperatures of the propellants to a suitable controller which is operably connected to valves between these manifolds and propellant storage tanks. When cryogenic propellant temperatures are sensed only a portion of the valves are opened to furnish propellants to some of the manifolds. When lower temperatures are sensed additional valves are opened to furnish propellants to more of the manifolds.

Abstract: There is disclosed an apparatus for and method of controlling a plant such as a turbojet engine wherein a model of the plant is created and plant performance request signals are applied to both the plant and the model of the plant. The performance of the model of the plant is used to modify the performance request signals supplied to the plant. Plant and plant model responses may also be compared to monitor degradation in performance of the plant. To adapt the scheme to nonlinear plants, the creation of a plurality of linear plant models each mimicking plant operation over a different limited portion of the toal range of plant performance is employed and there is a plant model for each of several different plant operation points about each of which, the actual plant operation is approximately linear.

Abstract: A method and apparatus 10 for self-regulating the mass flow rate of a fluid comprises a housing 12 defining an upper chamber 14 and a lower chamber 16, a sealed and pressurized bellows 30 contained within the upper chamber 14, spring bias means 58 contained within the lower chamber 16 and a poppet 32 with a poppet head 40 for defining a fluid passageway 44.Supply fluid flows through inlet 56, lower chamber 16, fluid passageway 44 and into upper chamber 14. Once in upper chamber 14 the fluid is in heat exchange communication with a fluid in sealed bellows 30. When the mass flow rate of the fluid flowing through upper chamber 14 is effected by either in pressure or a change in temperature, a corresponding change takes place in the fluid in the bellows 30 thus repositioning the poppet head 40 accordingly to redefine the area of passageway 44. Fluid leaves the apparatus through exits 28.

Abstract: The utilization of metal hydride and an acidic reagent for the accelerating of masses, and propulsion devices for applying such materials. The materials are of a saline complex metal hydride and acidic reagent which is sprayed onto the metal hydride which is offered in a lumpy consistency, for the pulse-like generation of expanding reaction gas bubbles in a constructive or dynamically, yieldably dammed chamber. Also provided is a propulsion device for the application of the above materials, with a support for a saline complex metal hydride of a lumpy consistency in a yieldably dammed expansion chamber in which an injection nozzle for the reagent is directed towards a surface portion of the metal hydride.

Abstract: A method is provided for controlling the utilization of a fluid bipropellant including two respective constituents separately in respective tanks aboard a spacecraft for consumption by a spacecraft rocket engine, comprising the steps of actuating the rocket engine; during the actuation of the rocket engine, providing a flow of bipropellant constituents to the rocket engine in a first proportion; after the actuation of the rocket engine, measuring the amount of at least one bipropellant constituent in a bipropellant tank containing the constituent; adjusting a pressure level within at least one bipropellant tank relative to a pressure level within another bipropellant tank based upon the amount of said at least one bipropellant constituent in a bipropellant tank containing said at least one constituent; after the measuring step, actuating the rocket engine; during the actuation of the rocket engine after the measuring step, providing a flow of bipropellant constituents to the rocket engine in a second proporti

Abstract: The process comprises the comparison of the flows of fuel and oxidizer measured at the outlet of the turbopumps with a set value corresponding to the optimum value of the mixture. The set value is periodically modified in dependence on the volumes of fuel and oxidizer remaining in the tanks. The control system for carrying out the process comprises a control valve controlled by an electronic device processing the measurements and disposed in a fuel or oxidizer supply pipe of a driving gas generator for the turbine of a turbopump or in the gas supply pipe of a turbopump, or disposed between the two pipes supplying the same fuel or oxidizer to two driving gas generators, or between the gas outlets of the two turbines of the turbopumps.

Abstract: A plug (24) has a fixed portion (26, 28) and a sliding portion (34). A tapered outer surface of portion (34) is adjacent to the throat (22) of a jet engine nozzle (20). Portion (34) includes two expansible chambers (36, 38) which are both defined by wall portions of both the fixed and the sliding portions of plug (24). Chambers (36, 38) are separated by a bulkhead (48) carried by said fixed portion. Chambers (36, 38) are supplied with fluid pressure which acts against a pressure surface (40, 42) to move sliding portion (34) axially forwardly and aftwardly, respectively, to in turn vary the throat area. Bleed air from the engine compressor supplies the fluid pressure. Fluid pressure is continuously supplied to chamber (38). Controlled leakage out of chamber (36) provides gradual slowing of movement of portion (34) in either direction. Control apparatus is located remotely from the nozzle (20).

Abstract: The process comprises the comparison of the flows of fuel and oxidizer measured at the outlet of the turbopumps with a set value corresponding to the optimum value of the mixture. The set value is periodically modified in dependence on the volumes of fuel and oxidizer remaining in the tanks. The control system for carrying out the process comprises a control valve controlled by an electronic device processing the measurements and disposed in a fuel or oxidizer supply pipe of a driving gas generator for the turbine of a turbopump or in the gas supply pipe of a turbopump, or disposed between the two pipes supplying the same fuel or oxidizer two driving gas generators, or between the gas outlets of the two turbines of the turbopumps.

Abstract: 1.

Abstract: Upon a landing approach, the normal compressor stator schedule of a fan speed controlled turbofan engine is temporarily varied to substantially close the stators to thereby increase the fuel flow and compressor speed in order to maintain fan speed and thrust. This running of the compressor at an off-design speed substantially reduces the time required to subsequently advance the engine speed to the takeoff thrust level by advancing the throttle and opening the compressor stators.