| Entries |
| Document | Title | Date |
|---|
| 20080264034 | SYSTEM AND METHOD FOR CONTROLLING THE TEMPERATURE AND INFRARED SIGNATURE OF AN ENGINE - A system and method for cooling at least a portion of an engine are provided. The engine is cooled using a fuel, such as a high heat sink fuel, that is subsequently used for combustion in the engine. The fuel can be used to cool one or more of the gases and/or components in the engine, thereby cooling the engine including an exhaust nozzle. For example, the fuel can be circulated through one or more heat exchanging devices that are disposed inside or outside a passage of the engine, and the fuel can absorb thermal energy from the engine or air that flows in the engine passage. In any case, the cooling of the engine can result in a reduction to the infrared signature of the engine. | 10-30-2008 |
| 20080282668 | Ether propeller - A propulsion system including an electromagnetic device for reaction with a medium; said propulsion system comprising:
| 11-20-2008 |
| 20080289317 | Control Method For Actuating a Thrust Reverser - The invention relates to a control method for opening or closing a turbojet engine thrust reverser ( | 11-27-2008 |
| 20090031699 | DRIVE LINE TORQUE PERTURBATION FOR PTO MODE SHIFTING - A method of controlling a machine drive having a driveline PTO establishes a driveline torque perturbation via conversion of internal inertia through two transmission neutral conditions. In an example, when a request is received, e.g., from an operator, to shift the driveline PTO from a first mode to a second mode, the transmission is automatically modulated between its first neutral condition and its second neutral condition while the driveline PTO is shifted from the first mode to the second mode, thus minimizing torque lock and facilitating mode changes. | 02-05-2009 |
| 20090071120 | Combined cycle integrated combustor and nozzle system - An engine that operates and produces the entire required vehicle thrust below Mach 4 is useful for a Hypersonic combined cycle vehicle by saving vehicle and engine development costs. One such engine is a combined cycle engine having both a booster and a dual mode ramjet (DMRJ). The booster and the DMRJ are integrated to provide effective thrust from Mach 0 to in excess of Mach 4. As the booster accelerates the vehicle from Mach 0 to in excess of Mach 4, from Mach 0 to about Mach 2 incoming air delivered to the DMRJ is accelerated by primary ejector thrusters that may receive oxidizer from either on-board oxidizer tanks or from turbine compressor discharge air. As the TBCC further accelerates the vehicle from about Mach 0 to in excess of Mach 4 exhaust from the turbine and exhaust from the DMRJ are combined in a common nozzle disposed downstream of a combustor portion of said DMRJ functioning as an aerodynamic choke. | 03-19-2009 |
| 20090133381 | Apparatus and method for testing performance of a material for use in a jet engine - An apparatus for testing failure of a material used in a jet engine, and more particularly to an apparatus that uses one or more model jet engine components made from a material used in a full-size jet engine and desired to be tested. The apparatus permits easy removal and disassembly of a jet engine mounted thereon as well as real-time measurements of run-time parameters. The methods and apparatus provide for predicting and analysing failure by a number of fatigue-related mechanisms including creep, fatigue, crack growth, foreign object damage, fretting, erosion, and stress corrosion. | 05-28-2009 |
| 20090235637 | SYSTEM, METHOD AND APPARATUS FOR LEAN COMBUSTION WITH PLASMA FROM AN ELECTRICAL ARC - The present invention provides a plasma arc torch that can be used for lean combustion. The plasma arc torch includes a cylindrical vessel, an electrode housing connected to the first end of the cylindrical vessel such that a first electrode is (a) aligned with a longitudinal axis of the cylindrical vessel, (b) extends into the cylindrical vessel, and (c) can be moved along the longitudinal axis, a linear actuator connected to the first electrode to adjust a position of the first electrode, a hollow electrode nozzle connected to the second end of the cylindrical vessel such that the center line of the hollow electrode nozzle is aligned with the longitudinal axis of the cylindrical vessel, and wherein the tangential inlet and the tangential outlet create a vortex within the cylindrical vessel, and the first electrode and the hollow electrode nozzle crate a plasma the discharges through the hollow electrode nozzle. | 09-24-2009 |
| 20090235638 | VARIABLE AREA NOZZLE ASSISTED GAS TURBINE ENGINE RESTARTING - A turbofan engine starting system includes a core nacelle housing ( | 09-24-2009 |
| 20090260343 | SOLID PROPELLANT MANAGEMENT CONTROL SYSTEM AND METHOD - Systems and methods of controlling solid propellant burn rate, propellant gas pressure, propellant gas pressure pulse shape, and propellant gas flow rate, rely on the position of a throttling valve. A throttling valve that is movable to a control position is disposed downstream of, and in fluid communication with, a solid propellant gas generator, and in parallel with a plurality of reaction control valves. The solid propellant in the solid propellant gas generator is ignited, to thereby generate propellant gas. The throttling valve is moved to a control position to attain a desired solid propellant burn rate, propellant gas pressure, and/or propellant gas pressure pulse shape. | 10-22-2009 |
| 20090266049 | METHODS AND SYSTEMS FOR PROPELLING A VEHICLE - Methods and systems are provided for propelling a vehicle. In an embodiment, by way of example only, a method includes flowing a decomposed hydroxylammonium nitrite (HAN)-based propellant into a chamber, introducing an aspirated non-polar fuel into the chamber, and combusting the decomposed HAN-based propellant and the aspirated non-polar fuel to produce an exhaust gas. | 10-29-2009 |
| 20090288386 | TRI-BODY VARIABLE AREA FAN NOZZLE AND THRUST REVERSER - A gas turbine engine system includes a fan ( | 11-26-2009 |
| 20090288387 | GAS TURBINE ENGINE BIFURCATION LOCATED IN A FAN VARIABLE AREA NOZZLE - A turbofan engine pylon structure having a fan variable area nozzle defined by a variable area flow system between a pylon intake and a pylon exhaust to selectively adjust a bypass flow. The variable area flow system changes the physical area and geometry of a fan nozzle exit area to manipulate the bypass flow by opening and closing an additional flow area of the variable area flow system. | 11-26-2009 |
| 20090293448 | Simplified thrust chamber recirculating cooling system - In some aspects a propulsion system includes a thrust chamber having a gap between an inner shell and an outer shell, the inner shell and the outer shell being attached together to form the thrust chamber. The rocket engine also includes a recirculating cooling system operably coupled to the gap in at least two locations and operable to recirculate a convective coolant through the gap. | 12-03-2009 |
| 20090313967 | POLYOXYMETHYLENE AS STRUCTURAL SUPPORT MEMBER AND PROPELLANT - A vehicle includes at least one polyoxymethylene structural support member. The polyoxymethylene structural support member includes a polyoxymethylene component that is a propellant that provides thrust to the vehicle upon pyrolysis or combustion of the polyoxymethylene component of the product of pyrolysis of the polyoxymethylene component. The vehicle can be a satellite or other type of spacecraft. | 12-24-2009 |
| 20090320442 | DUAL MODE PROPULSION SYSTEM - A system or method of propelling a vehicle with a multimodal propulsion system is provided. This multimodal propulsion system includes a primary inlet, a high-pressure gas generator, a gas generator exhaust system, a secondary fluid inlet, and a propulsion exhaust system. The primary inlet receives an incoming fluid flow that is provided to the high-pressure gas generator. The high-pressure gas generator coupled to the primary inlet produces a high-pressure exhaust from the incoming fluid flow. The high-pressure gas generator exhaust and secondary inlet couple to the propulsive exhaust system which mixes the two flow streams. The secondary flow may be gaseous or liquid fluid flow such as air flow for airborne flight and water for waterborne operation. The mixed fluid flow is expelled by the propulsive exhaust system that can propel the vehicle. | 12-31-2009 |
| 20100005777 | DUAL FUNCTION CASCADE INTEGRATED VARIABLE AREA FAN NOZZLE AND THRUST REVERSER - A gas turbine engine system includes a nozzle having a plurality of positions for altering a discharge flow received through the nozzle from a gas turbine engine fan bypass passage. The nozzle is integrated with a thrust reverser having a stowed position and a deployed position to divert the discharge flow and generate a reverse thrust force. At least one actuator is coupled with the nozzle and the thrust reverser to selectively move the nozzle between the plurality of positions and to move the thrust reverser between the stowed position and the deployed position. | 01-14-2010 |
| 20100011740 | TURBOFAN ENGINE OPERATION CONTROL | 01-21-2010 |
| 20100024386 | Gas-Generator Augmented Expander Cycle Rocket Engine - An augmented expander cycle rocket engine includes first and second turbopumps for respectively pumping fuel and oxidizer. A gas-generator receives a first portion of fuel output from the first turbopump and a first portion of oxidizer output from the second turbopump to ignite and discharge heated gas. A heat exchanger close-coupled to the gas-generator receives in a first conduit the discharged heated gas, and transfers heat to an adjacent second conduit carrying fuel exiting the cooling passages of a primary combustion chamber. Heat is transferred to the fuel passing through the cooling passages. The heated fuel enters the second conduit of the heat exchanger to absorb more heat from the first conduit, and then flows to drive a turbine of one or both of the turbopumps. The arrangement prevents the turbopumps exposure to combusted gas that could freeze in the turbomachinery and cause catastrophic failure upon attempted engine restart. | 02-04-2010 |
| 20100043389 | LOW SHOCK STRENGTH PROPULSION SYSTEM - Embodiments of the invention relate to a supersonic nacelle design employing a bypass flow path internal to the nacelle and around the engine. By shaping the nacelle, embodiments of the invention may function to reduce sonic boom strength, cowl drag, and/or airframe interference drag. Embodiments of the invention may also function to improve total pressure recovery and/or total thrust of the primary flow path through the engine. | 02-25-2010 |
| 20100043390 | CONTROLLING ICE BUILDUP ON AIRCRAFT ENGINE AND NACELLE STATIC AND ROTATING COMPONENTS - A turbofan engine deicing system includes a core nacelle ( | 02-25-2010 |
| 20100050594 | METHOD FOR REDUCING THE VIBRATION LEVELS OF A PROPFAN OF CONTRAROTATING BLADED DISKS OF A TURBINE ENGINE - The present invention relates to a method for reducing the vibration levels likely to occur, in a turbine engine comprising a first and a second bladed disk forming a propfan of contrarotating disks, when the two disks are traversed by a gaseous fluid, because of the turbulence of aerodynamic origin generated by the second bladed disk on the first bladed disk. The method comprises the following steps during the design of said two bladed disks: an initial configuration of the blades is defined, the synchronous forced response is calculated on the first bladed disk as a function of the harmonic excitation force generated by the second bladed disk expressed in the form of a linear function of the generalized aerodynamic force for the mode in question; for stacked sections of one of the two disks, a tangential geometric offset value θ of the individual aerodynamic profile is determined so as to reduce the term corresponding to the generalized aerodynamic force. The combination of the individual profiles on the sections with the tangential offsets therefore defines a new configuration of the blades of said one of the two disks which is applied to the blades of said one of the two disks. | 03-04-2010 |
| 20100058735 | OPERATIONAL LINE MANAGEMENT OF LOW PRESSURE COMPRESSOR IN A TURBOFAN ENGINE - A turbofan engine control system is provided for managing a low pressure compressor operating line. The engine includes a low spool having a low pressure compressor housed in a core nacelle. A turbofan is coupled to the low spool. A fan nacelle surrounds the turbofan and core nacelle and provides a bypass flow path having a nozzle exit area. A controller is programmed to effectively change the nozzle exit area in response to an undesired low pressure compressor stability margin which can result in a stall or surge condition. In one example, the physical nozzle exit area is decreased at the undesired stability condition occurring during engine deceleration. A low pressure compressor pressure ratio, low spool speed and throttle position are monitored to determine the undesired stability margin. | 03-11-2010 |
| 20100089028 | CORRUGATED CORE COWL FOR A GAS TURBINE ENGINE - An example core nacelle for a gas turbine engine includes a core cowl positioned adjacent to an inner duct boundary of a fan bypass passage having an associated cross-sectional area. The core cowl includes at least one groove that is selectively exposed to change the cross-sectional area. | 04-15-2010 |
| 20100101208 | Systems and Methods Involving Reduced Thermo-Acoustic Coupling of Gas Turbine Engine Augmentors - Systems and methods reducing thermo-acoustic coupling of a gas turbine engine augmentors are provided. In this regard, a representative method includes: determining acoustic resonances and heat release phase relationships associated with the augmentor; and determining relative axial positions of a fuel injector and a flame holder of the augmentor that result in reduced thermo-acoustic coupling of the augmentor during operation. | 04-29-2010 |
| 20100115914 | TRANSLATING CORE COWL HAVING AERODYNAMIC FLAP SECTIONS - An example core nacelle for a gas turbine engine includes a core cowl positioned adjacent an inner duct boundary of a fan bypass passage having an associated discharge airflow cross-sectional area. The core cowl includes at least one translating section and at least one flap section. The translating section of the core cowl is selectively moveable to vary the discharge airflow cross-sectional area. | 05-13-2010 |
| 20100139239 | GAS THRUSTER - A nozzle arrangement for use in a gas thruster is presented. At least one heater micro structure ( | 06-10-2010 |
| 20100139240 | BLADDER TYPE VARIABLE AREA FAN NOZZLE - A variable fan nozzle for use in a gas turbine engine includes a nozzle section, such as an inflatable bladder, associated with a fan bypass passage for conveying a bypass airflow. The nozzle section has an internal fluid pressure that is selectively variable to influence the bypass airflow. | 06-10-2010 |
| 20100146932 | GAS TURBINE ENGINE VARIABLE AREA EXHAUST NOZZLE - A gas turbine engine variable area exhaust nozzle has a plurality of first and second movable members arranged circumferentially around a fan casing extending in downstream. The first and second movable members alternate around the variable area nozzle. During cruise conditions the first and second movable members are in un-actuated positions. At take off conditions the first movable members are in an actuated position and the second movable members are in the un-actuated position. At top of climb conditions the first movable members are in the un-actuated position and the second movable members are in an actuated position. The first movable members move radially outwardly from the un-actuated position to the actuated position and the second movable members move radially inwardly from the un-actuated position to the actuated position. | 06-17-2010 |
| 20100162679 | GAS TURBINE ENGINE WITH EJECTOR - The present inventions include a boundary layer ejector fluidically connecting boundary layer bleed slots from an external surface of an aircraft to reduce aircraft/nacelle/pylon drag, reduce jet noise and decrease thrust specific fuel consumption. In one embodiment a boundary layer withdrawn through the boundary layer bleed slots is entrained with an exhaust flow of a gas turbine engine. In another embodiment a boundary layer withdrawn through the boundary layer bleed slots is entrained with a flow stream internal to the gas turbine engine, such as a fan stream of a turbofan. Members can be provided near an outlet of a passageway conveying the withdrawn boundary layer air to locally reduce the pressure of the fluid in which the withdrawn boundary layer air is to be entrained. A lobed mixer can be used in some embodiments to effect mixing between the boundary layer and a primary fluid of the ejector. | 07-01-2010 |
| 20100162680 | GAS TURBINE ENGINE WITH EJECTOR - The present inventions include a boundary layer ejector fluidically connecting boundary layer bleed slots from an external surface of an aircraft to reduce aircraft/nacelle/pylon drag, reduce jet noise and decrease thrust specific fuel consumption. In one embodiment a boundary layer withdrawn through the boundary layer bleed slots is entrained with an exhaust flow of a gas turbine engine. In another embodiment a boundary layer withdrawn through the boundary layer bleed slots is entrained with a flow stream internal to the gas turbine engine, such as a fan stream of a turbofan. Members can be provided near an outlet of a passageway conveying the withdrawn boundary layer air to locally reduce the pressure of the fluid in which the withdrawn boundary layer air is to be entrained. A lobed mixer can be used in some embodiments to effect mixing between the boundary layer and a primary fluid of the ejector. | 07-01-2010 |
| 20100180570 | Reactionless Motion (RM) - The Invention's Project of Reactionless Motion (RM) reveal a mechanically intrinsic motion technic, without motion opposites' reaction, or mass expelling. The RM invention consists in two circular and different reference frames of motion, under the influence of simultaneous attraction and repulsive forces between them. The consequence of said force's influence is a reaction less thrust in the center of each reference frame. Another effects of this invention are non typical observations in terms of propulsion technics, Newton's Third Law of Motion, space-time impulse response, kinetic energy, velocity and rotational motion. | 07-22-2010 |
| 20100180571 | MODULATING FLOW THROUGH GAS TURBINE ENGINE COOLING SYSTEM - A gas turbine engine system includes a fan bypass passage in a cooling passage having an inlet that receives a bleed flow from the fan bypass passage and an outlet that discharges the bleed flow into the fan bypass passage. A nozzle having a variable cross-sectional area controls an airflow within the fan bypass passage. The bleed flow outlet is placed such that moving the nozzle to change the variable cross-sectional area controls an amount of the bleed flow through the cooling passage. | 07-22-2010 |
| 20100192539 | METHODS OF CONTROLLING THRUST IN A ROCKET MOTOR - A propulsion thrust control system and method for controlling thrust in a rocket motor includes configuring valves of an energized rocket motor to an initial total valve area according to a total thrust command. The total thrust command is converted into a commanded propellant mass flow discharge rate. A varying total valve area is computed from an error between the commanded propellant mass flow discharge rate and a calculated propellant mass flow discharge rate. The valves are reconfigured according to a distribution of the varying total valve area. The propulsion system includes a pressure vessel with valves and a controller for regulating the valve area according to a propellant mass flow discharge rate from the pressure vessel. | 08-05-2010 |
| 20100223903 | VARIABLE PRESSURE RATIO COMPRESSOR - A compressor of a gas turbine engine may have a bypass that routes a compressed air flow from within the compressor and directs the compressed air flow to a combustor. The bypass may have an inlet positioned just ahead of a downstream stage of the compressor and an outlet positioned to route the compressed air flow from the bypass to a diffuser or directly to a combustor. A valve may be used within the bypass and may be located near the inlet, near the outlet, or both. The valve may have the form of an annular sleeve in some embodiments and may be actuated with an actuator. The various arrangements allow for a compressor having a variable compression ratio. | 09-09-2010 |
| 20100236216 | TURBOFAN ENGINE WITH VARIABLE AREA FAN NOZZLE AND LOW SPOOL GENERATOR FOR EMERGENCY POWER GENERATION AND METHOD FOR PROVIDING EMERGENCY POWER - A turbofan engine ( | 09-23-2010 |
| 20100242433 | METHOD FOR IMPROVING THE PERFORMANCE OF A BYPASS TURBOJET ENGINE - Method for improving the performance of a bypass turbojet engine. According to the invention, the area of the annular outlet orifice ( | 09-30-2010 |
| 20100257839 | Hydrocarbon-fueled rocket engine with endothermic fuel cooling - A rocket engine utilizes a kerosene-based fuel in a supercritical state which is catalytically converted to lighter hydrocarbons with heat from a thrust chamber assembly which operates as a heat exchanger. This process is facilitated by a fuel stabilization deoxygenator system which removes dissolved oxygen and/or by inerting the internal surfaces of the fuel-cooled combustion chamber wall passages by applying a zeolite-based catalyst coating to permit the fuel to be heated beyond normal temperature ranges. The supercritical kerosene-based fuel is passed through a turbine and injected into the combustion chamber to burn with the gaseous oxidizer. An increased mixing efficiency between the gaseous components results in an increase in combustion efficiency and increased stability of combustion. | 10-14-2010 |
| 20100269484 | SOLID PROPELLANT GAS CONTROL SYSTEM AND METHOD - Systems and methods of controlling solid propellant gas pressure and vehicle thrust are provided. Propellant gas pressure and a vehicle inertial characteristic are sensed. Propellant gas pressure commands and vehicle thrust commands are generated. A propellant gas pressure error is determined based on the propellant gas pressure commands and the sensed propellant gas pressure, and vehicle thrust error is determined based on the vehicle thrust commands and the sensed vehicle inertial characteristic. Reaction control valves are moved between closed and full-open positions based on the determined propellant gas pressure error and on the determined vehicle thrust error. The system and method allow the reaction control valves to operate at variable frequencies or at fixed frequencies. The system and method also allows propellant pressure to be commanded to follow a predetermined pressure profile or commanded to vary “on-the-fly.” | 10-28-2010 |
| 20100269485 | INTEGRATED VARIABLE AREA NOZZLE AND THRUST REVERSING MECHANISM - A variable area fan nozzle for use with a gas turbine engine system includes a nozzle having a cross-sectional area associated with a discharge air flow from a fan bypass passage. The nozzle is selectively moveable between multiple static positions for varying the cross-sectional area and multiple thrust reverse positions that are different from the static positions for reversing a direction of the discharge flow from the fan bypass passage to produce a thrust reversing force. | 10-28-2010 |
| 20100269486 | GAS TURBINE ENGINE HAVING SLIM-LINE NACELLE - A method of providing a slim-line nacelle for a gas turbine engine includes the steps of detecting a windmilling condition and increasing a discharge airflow area of a variable area fan nozzle in response to the detection of the windmilling condition. | 10-28-2010 |
| 20100275575 | METHOD AND SYSTEMS FOR CONTROLLING ENGINE THRUST USING VARIABLE TRIM - A method and systems for controlling a thrust output of a gas turbine engine are provided. The system includes a first sensor for measuring a first engine operating parameter, a second sensor for measuring a first engine condition parameter, and a processor programmed to determine an expected value of the first engine condition parameter and determine a first variance value using a difference between the expected value of the first engine condition parameter and the measured first engine condition parameter. The processor is further programmed to determine a trim value using the first variance value and a first engine operating parameter demand and to determine a modified operating parameter demand based on the nominal operating parameter demand and the determined trim value. The system also includes a controller coupled to the processor for controlling engine thrust based on the modified demand of the first engine operating parameter. | 11-04-2010 |
| 20100275576 | SYSTEM AND METHOD FOR MANEUVERING ROCKETS - A dual control system for a solid propellant propelled exoatmospheric kill-vehicle is disclosed. The system includes a primary solid propellant propulsion mechanism for controlling the propulsion of the kill-vehicle during a first phase of rocket flight, and a secondary solid propellant propulsion mechanism for maneuvering the kill-vehicle during a second phase of rocket flight. The primary propulsion mechanism controls the propulsion and direction of the kill-vehicle towards its target while the secondary propulsion mechanism may be used to provide lateral acceleration with a lower time-constant. | 11-04-2010 |
| 20100300065 | CLOSED-CYCLE ROCKET ENGINE ASSEMBLIES AND METHODS OF OPERATING SUCH ROCKET ENGINE ASSEMBLIES - Closed-cycle rocket engine assemblies including a combustor assembly, a combustor jacket, a turbine, a first pump and a mixing chamber are disclosed. The combustor jacket facilitates the transfer of heat from the combustor assembly into a fluid and the turbine is driven by a heated fluid from the combustor jacket. The mixing chamber may include a first inlet to receive a fluid from the turbine, a second inlet to receive a fluid from a first reactant reservoir, and an outlet to deliver a fluid to the first pump. Additionally, the first pump may be coupled to and powered by the turbine and the first pump may be configured to deliver at least a portion of the fluid from the mixing chamber into the combustion chamber of the combustor assembly. Related methods of operating such rocket engine assemblies are also disclosed. | 12-02-2010 |
| 20100313544 | PROPULSION SYSTEM WITH CANTED MULTINOZZLE GRID - A propulsion system includes a canted multinozzle plate, which has a multitude of small nozzles angled (not perpendicular) to major surfaces of the multinozzle grid plate. The multinozzle plate may be a cylindrical section or plate, and the multitude of nozzles may be substantially axisymmetric about the cylindrical plate. The propulsion system includes a pressurized gas source which may be placed either forward or aft of the multinozzle grid plate. The propulsion system may have a conical insert, an internal flow separator cone, to aid in changing directions of flow from the pressurized gas source, to divert the flow through the multiple nozzles. | 12-16-2010 |
| 20100313545 | GAS TURBINE ENGINE NOZZLE CONFIGURATIONS - In one embodiment, a gas turbine engine exhaust nozzle comprises a housing having a length which extends along a central longitudinal axis and comprising an interior surface and an exterior surface, and a row of chevrons extending from an aft end of the housing, the chevrons having a root region and a tip, wherein at least a portion of at least one of the interior surface or the exterior surface is scalloped proximate the root region of a chevron. Other embodiments may be described. | 12-16-2010 |
| 20110005192 | COOLING SYSTEM FOR AN AIRCRAFT, AIRCRAFT COMPRISING THE COOLING SYSTEM AND COOLING METHOD - A cooling system for an aircraft includes a cooling circuit for transporting a coolant, with the cooling circuit running to a component in the aircraft that is heated up during flying in order to take up heat from the heated-up component via the coolant. The cooling circuit runs to a jet engine for propulsion of the aircraft in order to release heat to a flow of gas in the jet engine via the coolant. | 01-13-2011 |
| 20110056183 | ULTRA-EFFICIENT PROPULSOR WITH AN AUGMENTOR FAN CIRCUMSCRIBING A TURBOFAN - An ultra-efficient “green” aircraft propulsor utilizing an augmentor fan is disclosed. A balanced design is provided combining a fuel efficient and low-noise high bypass ratio augmentor fan and a low-noise shrouded high bypass ratio turbofan. Three mass flow streams are utilized to reduce propulsor specific fuel consumption and increase performance relative to conventional turbofans. Methods are provided for optimization of fuel efficiency, power, and noise by varying mass flow ratios of the three mass flow streams. Methods are also provided for integration of external propellers into turbofan machinery. | 03-10-2011 |
| 20110088368 | THERMALLY-INTEGRATED FLUID STORAGE AND PRESSURIZATION SYSTEM - A propulsion system may be operated by determining pressure in a cryogenic liquid tank storing a fluid and cooling the cryogenic liquid tank in response to determining that the pressure is greater than a predetermined value. The cryogenic liquid tank may be pressurized by admitting a gaseous form of the fluid into the cryogenic liquid tank in response to determining that the pressure in the cryogenic liquid tank is less than a predetermined value. | 04-21-2011 |
| 20110088369 | METHOD FOR SYNCHRONIZING THE ACTUATORS OF A MOVABLE THRUST REVERSER COWL - The invention relates to a method for controlling a plurality of actuators ( | 04-21-2011 |
| 20110154804 | EXHAUST FOR A GAS TURBINE ENGINE - A gas turbine engine is provided having an offtake passage that in one form is capable of extracting a bypass flow from the engine. The airflow traversing the offtake passage is introduced to an exhaust flow of the gas turbine engine through an offtake outlet. The offtake outlet includes an airflow member that is moveable. A nozzle is also provided for exhaust from the gas turbine engine. In one form the nozzle includes moveable duct members. Flows exiting the offtake outlet and the nozzle can be combined after passing the airflow member and the moveable duct members, respectively. | 06-30-2011 |
| 20110167784 | METHOD OF OPERATING A CONVERTIBLE FAN ENGINE - A method of operating a gas turbine engine is disclosed having the steps of directing a flow of air to a front fan stage at a selected fan flow rate, fan speed and front fan stage pressure ratio corresponding to a selected first operating power level, pressurizing a portion of the flow from the front fan stage in a aft fan rotor to a first tip pressure ratio to generate a first overall fan pressure ratio, selecting a second operating power level that is lower than the first operating power level and reducing the flow in the aft fan rotor and pressurizing to a second tip pressure ratio to generate a second overall pressure ratio that is substantially lower than the first overall fan pressure ratio while the flow rate in the front fan stage is held substantially constant. | 07-14-2011 |
| 20110167785 | INTERNAL MIXING OF A PORTION OF FAN EXHAUST FLOW AND FULL CORE EXHAUST FLOW IN AIRCRAFT TURBOFAN ENGINES - An airborne mobile platform that has at least one turbofan engine assembly having a fan driven by a core engine, a short nacelle around the fan and a forward portion of the core engine, and a fan exhaust duct through the nacelle. A mixer duct shell is disposed substantially coaxial with and extending forwardly into the fan exhaust duct to provide a mixer duct between the mixer duct shell and a core engine shroud of the core engine. At least a portion of the mixer duct shell has a honeycomb core structure having an inner surface and an outer surface, with an acoustic lining on one of the inner or outer surfaces. The acoustic lining attenuates sound emanating from the turbofan engine assembly. | 07-14-2011 |
| 20110167786 | INTERNAL MIXING OF A PORTION OF FAN EXHAUST FLOW AND FULL CORE EXHAUST FLOW IN AIRCRAFT TURBOFAN ENGINES - An integrated, single piece mixer-center body ventilation apparatus is disclosed for use with a turbofan jet engine. The apparatus may incorporate a circumferential forward body portion adapted to be coupled to an aft end of a core engine turbine case of the jet engine, and a center body tube portion integrally formed with the forward body portion and having an axially opening vent exit. The forward body portion may have a plurality of inner mixer flow paths in communication with scalloped projecting portions. The inner mixer flow paths direct a pressurized core exhaust flow through the mixer device and mix the pressurized core exhaust flow with a portion of a pressurized fan exhaust flow, to thus significantly cool the pressurized core exhaust flow. | 07-14-2011 |
| 20110167787 | PULSE JET ENGINE - The present invention relates to pulse jet engines. More specifically, the present invention concerns a pulse jet engine which uses fluidic valving rather than unreliable mechanical valves, incorporates two combustion chambers with approximately the same but approximately 180 degree out-of-phase resonance frequencies to reduce noise through destructive interference of sound waves, and controls fuel injection and ignition within the two combustion chambers to increase efficiency by achieving more rapid and complete combustion. | 07-14-2011 |
| 20110167788 | THRUST VECTORING APPARATUS FOR A JET ENGINE, CORRESPONDING JET ENGINE, THRUST VECTORING METHOD AND UPGRADING METHOD FOR A JET ENGINE - The thrust vectoring apparatus comprises: a housing defining a primary outlet for emitting the primary jet; Coanda surfaces extending from opposing regions of said housing, and radially spaced from the primary outlet such that a step is defined between each Coanda surface and the primary outlet; ducts leading from a fluid source to secondary outlets; and flow control means operable to control the mass flow through the secondary outlets. When the jet engine operates to exhaust a primary jet through the primary outlet, low pressure regions are formed in the vicinity of the steps. Each secondary outlet is located adjacent one of the Coanda surfaces so as to emit a secondary flow into a low pressure region. On activation of the secondary flow by the flow control means, the primary jet is entrained by the Coanda surface opposing the Coanda surface adjacent said the secondary outlet from which the secondary flow has been emitted. Method of vectoring the thrust or of upgrading existing jet engines with the thrust vectoring apparatus and jet engines comprising the thrust vectoring apparatus are disclosed | 07-14-2011 |
| 20110179766 | HEAT RECOVERY SYSTEM - An improved turbofan engine having an engine shaft and a centerbody inside of an exhaust nozzle, the turbofan engine having a fluid-filled conduit located about the centerbody and a heat recovery apparatus operatively connected to the fluid-filled conduit. | 07-28-2011 |
| 20110203253 | ADVANCED FUEL COMPOSITIONS FROM RENEWABLE SOURCES, AND RELATED METHODS FOR MAKING AND USING THE FUEL - A method for preparing a fuel composition is described, and includes the step of preparing a bio-derived fuel component that contains a mixture of iso-saturated alkanes and normal-saturated alkanes. The method further includes the step of determining if the ratio of iso-saturated alkanes to normal-saturated alkanes is at least about 2.0. If that requirement is met, the bio-derived fuel component is usually combined with a petroleum-derived component, resulting in the fuel composition. Related compositions are also described, in which the weight ratio of iso-saturated alkanes to normal-saturated alkanes is at least about 2.0; and the composition has a freeze point less than about −50° C. | 08-25-2011 |
| 20110214409 | Combustion Turbine in which Combustion is Intermittent - The subject of the present invention is a combustion turbine ( | 09-08-2011 |
| 20110219742 | SUPERSONIC COMBUSTOR ROCKET NOZZLE - A supersonic combustor as a component of a rocket nozzle offers improved utilization of available chemical energy that may be released from combustion gasses flowing through the rocket nozzle. A subsonic combustor sub-sonically accelerates an exothermically reacting combustion gas up to a nozzle throat. The supersonic combustor expands and super-sonically accelerates the exothermically reacting combustion gas beyond the nozzle throat. The dimensions of the supersonic combustor may be selected such that the supersonic combustor achieves a slow rate of cooling of the combustion gasses without creating shockwaves within the supersonic combustor. A supersonic discharge expands and super-sonically accelerates the now substantially non-reacting combustion gas through a supersonic discharge of the rocket nozzle. The momentum of the combustion gas leaving the supersonic discharge propels the rocket nozzle in the opposite direction due to the principle of conservation of momentum. | 09-15-2011 |
| 20110277446 | ROCKETS, METHODS OF ROCKET CONTROL AND METHODS OF ROCKET EVALUATION UTILIZING PRESSURE COMPENSATION - Rockets, rocket motors, methods of controlling a rocket and methods of evaluating a rocket design are disclosed. In some embodiments, a method of controlling a rocket may include measuring a combustion chamber pressure, calculating a logarithm of the measured combustion chamber pressure, and computing the difference between the logarithm of the measured combustion chamber pressure and the logarithm of a reference combustion chamber pressure value to generate an error signal. The method may further include filtering the error signal to generate a compensated signal in the logarithm domain, and exponentiating of the compensated signal in the logarithm domain to provide a compensated signal in the physical domain. | 11-17-2011 |
| 20110302905 | EYEBALL SEALS FOR GIMBALED ROCKET ENGINES, AND ASSOCIATED SYSTEMS AND METHODS - Eyeball seals for a gimbaled rocket engines, and associated systems and methods are disclosed. A system in accordance with a particular embodiment includes a rocket body, an engine carried by and movable relative to the rocket body, and a seal assembly. The seal assembly can include a sealing surface carried by one of the rocket body and the engine, and a seal element carried by the other of the rocket body and the engine. The seal element is in contact with the sealing surface. The seal assembly can further include a cylinder and a piston slideably received in the cylinder, with one of the piston and the cylinder carrying the seal element. The cylinder includes ports that are in fluid communication with a region external to the rocket body. Accordingly, pressures external to the rocket body can force the seal element and/or the sealing surface into contact with each other. | 12-15-2011 |
| 20110302906 | Laser Tractor Beam - There is provided a method of using a remote laser source to manipulate a space object having a target, comprising projecting a beam from the remote laser source, wherein the beam has a sufficient intensity and wavelength to cause ablation at a position on the target; imparting an impulse to the space object having the target; modifying at least one beam characteristic selected from the group consisting of intensity, wavelength and position on the target, wherein the position and/or orientation of the space object is altered relative to the remote laser source. | 12-15-2011 |
| 20110308231 | SYSTEM, METHOD AND APPARATUS FOR COOLING ROCKET MOTOR COMPONENTS USING A SATURATED LIQUID VAPOR COOLANT MIXTURE - A system and method of cooling a rocket motor component includes injecting a high pressure liquid coolant through an injector nozzle into a cooling chamber. The cooling chamber having a pressure lower than the high pressure liquid coolant. The liquid coolant flashes into a saturated liquid-vapor coolant mixture in the cooling chamber. The saturated liquid-vapor coolant mixture is at equilibrium at the lower pressure of the cooling chamber. Heat from the rocket motor component to be cooled is absorbed by the coolant. A portion of the liquid portion of the saturated liquid-vapor coolant mixture is converted into gas phase, the converted portion being less than 100% of the coolant. A portion of the coolant is released from the cooling chamber and the coolant in the cooling chamber is dynamically maintained at less than 100% gas phase of the coolant as the thrust and heat generated by the rocket motor varies. | 12-22-2011 |
| 20110308232 | Method and device to increase thrust and efficiency of jet engine - A method and device for adding special substances to the gaseous outflow of the jet engine in order to create a “virtual wall” of increased pressure zone behind the jet engine nozzle which can serve as a support for the jet engine gas outflow to push against, thus increasing the thrust power of the jet engine. This increase in acceleration power results in the accelerated movement of the jet engine equipped vehicle or higher fuel efficiency. A device that adds the special substances to the gaseous outflow is proposed. Characteristics of the special substances, which if added into the jet engine gaseous outflow may create a “virtual wall” of increased pressure zone behind the jet engine, are proposed. | 12-22-2011 |
| 20120011825 | GAS TURBINE ENGINE WITH NOISE ATTENUATING VARIABLE AREA FAN NOZZLE - A bypass gas turbine engine includes a variable area fan nozzle with a leading edge region that defines an increased airfoil leading edge radius. | 01-19-2012 |
| 20120096830 | TURBINE AND METHOD THEREOF - The present disclosure relates to an improved and more efficient turbine deployed in the generation of thermal energy. The improved turbine comprises aerodynamic blades and supersonic nozzles to generate impulse and reaction forces which results into higher efficiency. | 04-26-2012 |
| 20120124962 | METHOD OF VARYING A FAN DUCT THROAT AREA - A method of varying a fan duct throat area of a gas turbine engine may include non-pivotably moving a fan nozzle outwardly relative to a longitudinal axis of the gas turbine engine during axially aft translation of the fan nozzle without activating a thrust reverser. | 05-24-2012 |
| 20120124963 | METHOD OF VARYING A FAN DUCT NOZZLE THROAT AREA OF A GAS TURBINE ENGINE - A method of varying a fan duct nozzle throat area of a gas turbine engine includes pivoting a fan nozzle outwardly relative to a longitudinal axis of the gas turbine engine. The fan nozzle is configured to move axially non-contemporaneously with the pivoting of the fan nozzle. | 05-24-2012 |
| 20120124964 | GAS TURBINE ENGINE WITH IMPROVED FUEL EFFICIENCY - A turbofan engine includes a fan driven by a low pressure turbine through a gear reduction. The gear reduction has a gear ratio of greater than or equal to about 2.4. The low pressure turbine has an expansion ratio greater than or equal to about 5. The fan has a bypass ratio greater than or equal to about 8. In other features, a turbofan engine includes a variable geometry fan exit guide vane (FEGV) system having a multiple of circumferentially spaced radially extending fan exit guide vanes. Rotation of the fan exit guide vanes between a nominal position and a rotated position selectively changes a fan bypass flow path to permit efficient operation at various flight conditions. | 05-24-2012 |
| 20120131901 | SYSTEM AND METHOD FOR CONTROLLING A PULSE DETONATION ENGINE - In one embodiment, a pulse detonation engine (PDE) includes a controller configured to receive signals indicative of at least one of a desired operating parameter of the PDE and a measured internal parameter of the PDE, and to adjust at least one of a first fluid flow through the PDE and a second fluid flow through at least one of multiple pulse detonation tubes disposed within the PDE based on the signals. The PDE does not include a turbine or a mechanical compressor. | 05-31-2012 |
| 20120137653 | MULTI-STAGE ROCKET, DEPLOYABLE RACEWAY HARNESS ASSEMBLY AND METHODS FOR CONTROLLING STAGES THEREOF - A deployable raceway harness assembly for use with a multi-stage rocket includes a first cable bundle configured to extend across a second stage of a multi-stage rocket from a guidance unit to a first stage of the rocket. The deployable raceway harness assembly includes a deployable raceway cover configured for detachable coupling with the multi-stage rocket. The deployable raceway cover extends over at least the first cable bundle and a second cable bundle. The first cable bundle is fastened to the deployable raceway cover. The second cable bundle is configured to extend from the guidance unit to the second stage and is shorter than the first cable bundle. An in-flight deployment mechanism is configured to detach the deployable raceway cover and the first cable bundle extending across the second stage from the multi-stage rocket in-flight leaving the second cable bundle extending to the second stage in place. | 06-07-2012 |
| 20120137654 | Thrust Reverser and Variable Area Fan Nozzle Actuation System and Method - There is provided an actuation system for a gas turbine engine including a thrust reverser and a variable area fan nozzle. The system has a plurality of linear actuators each having a first outer piston concentric with a second inner piston. The first outer piston is operatively connected to a thrust reverser. The second inner piston is operatively connected to a variable area fan nozzle. The system further has a piston lock assembly for selectively locking the first outer piston to the second inner piston. The system further has a control system coupled to the plurality of linear actuators for operating the variable area fan nozzle between a stowed position and a deployed position. | 06-07-2012 |
| 20120137655 | VARIABLE AREA FAN NOZZLE THRUST REVERSER - A nozzle for use in a gas turbine engine includes nozzle doors coupled with a fan nacelle wherein the nozzle doors move in unison between a plurality of positions to influence a bypass airflow through a fan bypass passage. A linkage connects the nozzle doors and an actuator. A louver section coupled with the linkage moves in unison with the nozzle doors between a plurality of louver positions to direct a portion of the bypass airflow in a selected direction. | 06-07-2012 |
| 20120151897 | SYSTEM AND METHOD FOR OPERATING A THRUST REVERSER FOR A TURBOFAN PROPULSION SYSTEM - A thrust reverser assembly for use in a turbofan engine assembly. The engine assembly includes a core gas turbine engine and a core cowl which circumscribes the core gas turbine engine. A nacelle is positioned radially outward from the core cowl to define a fan nozzle duct between the core cowl and a portion of the nacelle. The nacelle includes a stationary cowl. The thrust reverser assembly includes a first translating cowl that is slidably coupled to the nacelle. The first translating cowl is positionable with respect to the stationary cowl. A second translating cowl is slidably coupled to the nacelle such that the first translating cowl is positioned between the stationary cowl and the second translating cowl. The second translating cowl is positionable with respect to the first translating cowl. A positioning assembly is coupled to the first translating cowl. An actuator assembly is operatively coupled to the second translating cowl for selectively moving the second translating cowl. The actuator assembly is configured to engage the positioning assembly to selectively move the first translating cowl. | 06-21-2012 |
| 20120151898 | SYSTEM AND METHOD OF COMBUSTION FOR SUSTAINING A CONTINUOUS DETONATION WAVE WITH TRANSIENT PLASMA - An annular combustion ( | 06-21-2012 |
| 20120151899 | MORPHING STRUCTURE AND METHOD - A method of controlling mixing of a flow exiting a downstream end of a primary nozzle associated with a jet engine. The method may involve coupling a shape memory alloy (SMA) element to a mixing structure disposed at the downstream edge of the primary nozzle. An electrical signal may be applied to the SMA element to heat the SMA element and induce a phase change in the SMA element. The phase change may cause an axial length of the SMA element to constrict, to cause movement of the mixing structure into a path of the flow exiting the primary nozzle. | 06-21-2012 |
| 20120159925 | TURBINE BASED COMBINED CYCLE ENGINE - An aircraft powerplant is disclosed that can be operated in at least three modes including as a gas turbine engine, as an engine having a ramburner, and as an engine having a forward compression combustor engine such as a ramjet and/or scramjet. An airflow valve is provided to direct air to a downstream portion of the aircraft engine and can be positioned as a function of the aircraft operating modes. The valve can be used to cocoon the gas turbine engine. | 06-28-2012 |
| 20120167549 | FLADE DISCHARGE IN 2-D EXHAUST NOZZLE - An aircraft gas turbine engine has a row of FLADE fan blades disposed radially outwardly of and drivingly connected to a fan in the engine's fan section. The FLADE fan blades extend across a FLADE duct circumscribing the fan section. A two dimensional air discharge passage is in fluid flow communication with the FLADE duct and with FLADE air upstream and downstream discharge slots in a divergent flap of a two dimensional exhaust nozzle. A valve fully closes the upstream slot when the downstream slot is fully opened and fully opens the upstream slot when the downstream slot is fully closed. The upstream and downstream slots may be located upstream and downstream respectfully of a nozzle discharge area in the nozzle. A sliding deck slides aft or down to open upstream slot and close downstream slot and slides forward or up to close upstream slot and open downstream slot. | 07-05-2012 |
| 20120174559 | SYSTEM AND METHOD FOR CONTROLLING A GAS TURBINE ENGINE AFTERBURNER - Methods and apparatus are provided for operating a gas turbine engine. In a first operational mode, the gas turbine engine generates thrust using the propulsion turbine and not the afterburner when it is commanded to generate a thrust between at least a first thrust magnitude and a second thrust magnitude, and generates thrust using the propulsion turbine and the afterburner when it is commanded to generate thrust greater than the second thrust magnitude. In a second operational mode, the gas turbine engine generates thrust using the propulsion turbine and the afterburner when it is commanded to generate a thrust greater than the first thrust magnitude. The steady state thrust-versus-throttle position response has a substantially constant linear slope that is set to be two times the similar slope of the first operational mode. Thrust transients of the propulsion turbine and the afterburner are substantially synchronizing when the gas turbine engine is operating in the second mode and generating thrust greater than first thrust magnitude. | 07-12-2012 |
| 20120180451 | Optimal Feedback Heat Energy Internal Combustion Engine And Its Applications - An internal combustion engine wherein a thermo potential heat flow in combustion is maximised by providing a feedback of an optimised amount of thermo potential heat flow that is modulated in the exhaust media, into the air intake, and a method of providing feedback comprises producing a shock wave of pulse of exhaust media and pulse of intake air on the opposite side of a high temperature sustainable wire screen modem thereby transferring the thermo potential heat energy flow from the exhaust media to the air intake. | 07-19-2012 |
| 20120192543 | EXHAUST NOZZLE FOR A BYPASS AIRPLANE TURBOJET HAVING A DEPLOYABLE SECONDARY COVER AND A RETRACTABLE CENTRAL BODY - The invention relates to an exhaust nozzle for a bypass airplane turbojet comprising an annular central body, an annular primary cover surrounding the central body to define a hot stream flow channel, and an annular secondary cover surrounding the primary cover to define a cold stream flow channel, each of the central body and the secondary cover comprising a stationary portion and a movable portion connected to a downstream end of the stationary portion, the movable portion of the central body being suitable for being retracted longitudinally upstream relative to the stationary portion, and the movable portion of the secondary cover being suitable for being deployed longitudinally downstream relative to the stationary portion. | 08-02-2012 |
| 20120198813 | THRUST CHAMBER AND ROCKET ENGINE SYSTEM - An engine system includes a thrust chamber that has a cooling channel. The cooling channel is adapted to provide sustained cracking conditions for a fluid at steady-state operating conditions. A turbine has an input in fluid communication with an output of the cooling channel. A pump is mechanically coupled with the turbine and is in fluid communication with the cooling channel. | 08-09-2012 |
| 20120198814 | USE OF HOT GASES AND DEVICES - A method of increasing internal combustion engine efficiency is based on using engine cooling air and exhaust gas by flowing this mixture into a convergent nozzle thus accelerating the gas mixture and eject it through nozzle exit, thus generating thrust in a desired direction which could push a land air or sea vehicle. Another option is to use the accelerated gas to drive a turbine that could add its torque to the engine or to drive electrical generator that produces electricity. | 08-09-2012 |
| 20120204534 | SYSTEM AND METHOD FOR DAMPING PRESSURE OSCILLATIONS WITHIN A PULSE DETONATION ENGINE - In one embodiment, a pulse detonation engine includes a resonator configured to fluidly couple to an air flow path upstream of a pulse detonation tube. The pulse detonation engine also includes a controller configured to receive signals indicative of an operating frequency of an air valve disposed at an upstream end of the pulse detonation tube, and to adjust a geometric configuration of the resonator in response to the signals. | 08-16-2012 |
| 20120227374 | INTEGRATED VEHICLE FLUIDS - A system and methods are provided for combining systems of an upper stage space launch vehicle for enhancing the operation of the space vehicle. Hydrogen and oxygen already on board as propellant for the upper stage rockets is also used for other upper stage functions to include propellant tank pressurization, attitude control, vehicle settling, and electrical requirements. Specifically, gases from the propellant tanks, instead of being dumped overboard, are used as fuel and oxidizer to power an internal combustion engine that produces mechanical power for driving other elements including a starter/generator for generation of electrical current, mechanical power for fluid pumps, and other uses. The exhaust gas from the internal combustion engine is also used directly in one or more vehicle settling thrusters. Accumulators which store the waste ullage gases are pressurized and provide pressurization control for the propellant tanks. The system is constructed in a modular configuration in which two redundant integrated fluid modules may be mounted to the vehicle, each of the modules capable of supporting the upper stage functions. | 09-13-2012 |
| 20120233979 | ROCKET MULTI-NOZZLE GRID ASSEMBLY AND METHODS FOR MAINTAINING PRESSURE AND THRUST PROFILES WITH THE SAME - A rocket multi-nozzle grid assembly includes an insulator grid, a plurality of refractory nozzle fittings and a grid support plate. The grid support plate braces the plurality of refractory nozzle fittings and the insulator grid. The insulator grid includes insulator orifices, each of the nozzle fittings includes a fitting orifice and the grid support plate includes a plurality of plate orifices. The rocket multi-nozzle grid assembly includes a plurality nozzles. Each nozzle includes a plate orifice aligned with one insulator orifice and one fitting orifice. Exhaust gases are directed through the plurality of nozzles. The multi-nozzle grid assembly substantially maintains its surface area throughout operation of a rocket motor. In one example, the surface area is maintained through ablation of the insulator grid. In another example, the surface area is maintained through cooling of the plurality of nozzle fittings with the ablated fragments passing through the nozzles. | 09-20-2012 |
| 20120240550 | Vehicle Propulsion System - A vehicle propulsion system having a manifold in fluid communication with a cross flow fan which adjusts to control an amount lift generated by a plurality of airfoils providing lift from air from the cross flow fan. | 09-27-2012 |
| 20120279196 | MICROFLUIDIC FLAME BARRIER - Propellants flow through specialized mechanical hardware that is designed for effective and safe ignition and sustained combustion of the propellants. By integrating a micro-fluidic porous media element between a propellant feed source and the combustion chamber, an effective and reliable propellant injector head may be implemented that is capable of withstanding transient combustion and detonation waves that commonly occur during an ignition event. The micro-fluidic porous media element is of specified porosity or porosity gradient selected to be appropriate for a given propellant. Additionally the propellant injector head design integrates a spark ignition mechanism that withstands extremely hot running conditions without noticeable spark mechanism degradation. | 11-08-2012 |
| 20120304619 | ENGINE FOR THRUST OR SHAFT OUTPUT AND CORRESPONDING OPERATING METHOD - An engine and a method of operating the engine are provided. The engine includes a gas turbine and fan that rotate together to provide an exhaust gas flow stream, which flows over a free turbine that is connected to a power take-off. The free turbine can extract energy from the exhaust gas flow stream and transfer it as shaft power to the power take-off and the amount of energy extracted by the free turbine is controlled by varying the pitch of the free turbine's blades and/or by varying the pitch or stator vanes of a stator upstream of the free turbine. The control over the amount of energy extracted by the free turbine allows the engine to be used to provide thrust from the gas turbine and fan or to provide shaft power at the power take-off, or a combination of thrust and shaft power. | 12-06-2012 |
| 20120317956 | Constant Volume Combustion Chamber - A constant volume combustion chamber, combustor, and method for constant volume combustion involve combusting a fuel in a chamber sealed by a pintle having a conical portion fitted into a conical nozzle throat and pulling the pintle away from the nozzle throat to allow combustion products to exhaust through a nozzle outlet. The shapes and surfaces of the pintle and nozzle throat provide for sealing the chamber at high pressures while resisting surface wear. Operational parameters for the combustor may be computer controlled in response to measured pressures and temperatures in the combustor. | 12-20-2012 |
| 20120317957 | AIRCRAFT POWERPLANT - A gas turbine engine system is disclosed which includes a core passage and a bypass passage which can be configured as a fan bypass duct or a third stream bypass duct. The core passage and bypass passage are routed to flow through a nozzle before exiting overboard an aircraft. The nozzle includes moveable members capable of changing a configuration of the nozzle. In one form the moveable members are capable of changing throat area for portions of the nozzle that receive working fluid from the core passage and the bypass passage. The bypass passage can include a branch. In one form the branch can include a heat exchanger. The bypass passage can also provide cooling to one or more portions of the nozzle, such as cooling to a deck of the nozzle. | 12-20-2012 |
| 20130025256 | Pulsed Detonation Engine - A pulsed detonation engine may include a detonation tube for receiving fuel and an oxidizer to be detonated therein, one or more fuel-oxidizer injectors for injecting the fuel and oxidizer into the detonation tube, one or more purge air injectors for injecting purge air into the detonation tube for purging the detonation tube, and an ignition for igniting the fuel and oxidizer in the detonation tube so as to initiate detonation thereof. The detonation tube has an upstream end, a downstream end, and an axially extended portion extending from the upstream end to the downstream end and having a perimeter. The fuel-oxidizer injectors and purge air injectors may be disposed at least along the axially extended portion. The ignition may include a plurality of igniters disposed at or near the perimeter of the axially extended portion, spaced about the perimeter, at or near the upstream end of the detonation tube. | 01-31-2013 |
| 20130061571 | Laser propelled flight vehicle - Disclosed are embodiments for producing thrust, and in particular thrust for the propulsion of a flight vehicle. The embodiments incorporate the on-board laser heating of a propellant to a plasma state for the production of thrust, and of energy being supplied to the on-board laser by remote power sources such as ground based, sea based, space based, or airborne pump lasers. | 03-14-2013 |
| 20130067884 | THRUST REVERSER FOR A GAS TURBINE ENGINE - A thrust reverser includes a slider movable along an actuator shaft. An inner linkage is mounted to the slider and the inner thrust reverser door and an outer linkage is mounted to the slider and the outer thrust reverser door. | 03-21-2013 |
| 20130074472 | INJECTOR HAVING MULTIPLE IMPINGEMENT LENGTHS - An injector includes an injector body that has at least an injection side surface. A plurality of passages extends within the injector body for injecting combustion fluids at the injection side surface. The plurality of passages include a first set of passages that has a first arrangement defining a first impingement point in space with regard to the injection side surface and a second, different set of passages that has a second arrangement defining a second impingement point in space with regard to the injection side surface. The first impingement point and the second impingement point are non-equidistant from the injection side surface | 03-28-2013 |
| 20130104521 | GAS TURBINE ENGINE WITH TWO-SPOOL FAN AND VARIABLE VANE TURBINE | 05-02-2013 |
| 20130104522 | GAS TURBINE ENGINE WITH VARIABLE PITCH FIRST STAGE FAN SECTION | 05-02-2013 |
| 20130125527 | REVERSIBLE FLOW DISCHARGE ORIFICE - A rocket engine fluid-flow system includes a pump fluidly interconnecting a fluid source to a combustion chamber. A nozzle is in fluid communication with the combustion chamber and includes coolant tubes fluidly arranged between the pump and the combustion chamber. An orifice has a throat and is fluidly arranged between the pump and the coolant tubes. The orifice has entrance and exit ramps arranged on either side of the throat. The exit ramp has an exit ramp surface with a divergent angle that is less than a right angle. The entrance ramp provides a smooth approach to the orifice throat. In one example, the exit ramp includes an exit ramp surface having a divergent angle of 20-60°. The exit ramp radius is less than twice the throat radius in one example. | 05-23-2013 |
| 20130186058 | GEARED TURBOMACHINE FAN AND COMPRESSOR ROTATION - A high-bypass ratio geared turbomachine comprises a compressor section of a high-bypass ratio geared turbomachine. The compressor section provides at least a low-pressure compressor and a high-pressure compressor, wherein a rotor of the low-pressure compressor rotates together with a rotor of a fan. | 07-25-2013 |
| 20130192195 | GAS TURBINE ENGINE WITH COMPRESSOR INLET GUIDE VANE POSITIONED FOR STARTING - A gas turbine engine includes a variable inlet guide vane positioned forwardly of a low pressure compressor. The angle of the inlet guide vane is controlled at startup to increase airflow into the compressor. This is particularly useful when the gas turbine engine is being restarted while an associated aircraft is in the air, and is relied upon to increase windmill speed of the compressor and turbine rotors. A method and variable inlet vane are also disclosed. | 08-01-2013 |
| 20130199153 | METHOD AND APPARATUS TO CONTROL PART-LOAD PERFORMANCE OF A TURBINE - A method of controlling the part-load performance of a turbine includes generating a bypass flow in the turbine by removing a portion of a compressed fluid from a compressor of the turbine, determining an operating load of the turbine, transmitting the bypass flow to a turbine section of the turbine; and selectively heating the bypass flow according to the determined operating load of the turbine. | 08-08-2013 |
| 20130199154 | RAMJET SUPERHEATER - The present invention provides a regenerative superheater system for an ejector ramjet engine. The invention includes a superheater in thermal communication with the combustion chamber of the ramjet engine. The superheater transfers thermal energy from combustion chamber to an ejectant which is then redirected upstream to the ramjet ejector. In one embodiment of the invention the temperature of the ejectant is modulated by a variable geometry cooler that controls the amount of thermal energy removed from the superheater system by ambient air. In an alternate embodiment of the invention, the temperature of the ejectant is modulated by a variable geometry superheater that controls the amount of thermal energy added to the superheater system through combustion gas. | 08-08-2013 |
| 20130213006 | ACTUATOR ASSEMBLY - An actuator assembly includes a rotatable first shaft, a second shaft, and a ball screw mechanism. The second shaft is telescopically received in and extensible relative to the first shaft. The second shaft is rotatable with the first shaft and is connected to and drives the ball screw mechanism. | 08-22-2013 |
| 20130219854 | AIR-COOLED OIL COOLER FOR TURBOFAN ENGINE - A turbofan gas turbine engine comprises a nacelle cowl and a core engine. A bypass duct is between an outer surface of a casing of the core engine, and an inner surface of the nacelle cowl. An air channel is in the nacelle cowl, an inlet and an outlet of the air channel being in an outer surface of the nacelle cowl. An oil cooler has at least one oil passage for oil circulation, the air cooler having a first heat exchange surface in the air channel exposed to air circulating in the air channel, the air channel having a second heat exchange surface in the bypass duct exposed to air circulating in the bypass duct. A method for cooling oil in a turbofan gas turbine engine is also provided. | 08-29-2013 |
| 20130219855 | Method of Using an Afterburner to Reduce High Velocity Jet Engine Noise - A system and method of reducing noise caused by jet engines is provided. The method comprises the steps of using the afterburner to heat the exhaust gas flow while simultaneously reducing power to the core engine. Together these two operations reduce the pressure of the exhaust gas in the nozzle area while holding the exhaust gas velocity constant, which maintains engine thrust while decreasing engine noise. The method may be supplemented by altering the location of the afterburner flames to create an inverted exhaust velocity profile, thereby decreasing engine noise even further. | 08-29-2013 |
| 20130232948 | Noise-Reducing Engine Nozzle System - A method and apparatus for reducing noise generated during operation of an engine. In one illustrative embodiment, a nozzle system comprises a first nozzle and a second nozzle at least partially surrounded by the first nozzle. An outer surface of an aft portion of the second nozzle has a shape configured such that a radial cross-section of the outer surface of the aft portion of the second nozzle has a curve that is different from at least one other curve for another radial cross-section of the outer surface of the aft portion of the second nozzle and such that an axial cross-section of the outer surface of the aft portion of the second nozzle has a wavy shape. | 09-12-2013 |
| 20130239544 | Distributed pressurization system - The present invention comprises a distributed system for providing warm, high pressure gas that is lighter weight, lower cost, and more reliable than comparable currant art pressurization systems. The invention employs Tridyne or another non-explosive but combustible pressurant mix flowing through a plurality of catalytic devices to provide heated pressurant to both the needed rocket system applications as well as to the plurality of pressurant storage bottles, thereby decreasing the mass of the unused pressurant at the end of mission. | 09-19-2013 |
| 20130247542 | SPACE LAUNCHER PROPULSION SYSTEM IMPLEMENTING A METHOD OF REGULATING PROPELLANT COMPONENT CONSUMPTION - A space launcher propulsion system includes: an engine fed by at least two propellant component tanks; a measurement mechanism to measure a quantity of propellant component actually consumed in each of the tanks, by sensors becoming uncovered; an estimation mechanism to estimate instantaneous flow rates of each of the propellant components on the basis of at least one operating parameter of the engine; and a correction mechanism correcting the way the estimation mechanism estimates on the basis of the measured consumed quantities and of the estimated instantaneous flow rates. | 09-26-2013 |
| 20130269312 | DEVICE FOR CONTROLLING A VARIABLE SECTION NOZZLE OF AN AIRCRAFT - A control device for controlling a variable section nozzle of an aircraft power plant, the variable section nozzle including one or several movable parts capable of modifying the nozzle section and connected by a mechanical transmission chain to an actuator. The control device includes a system for regulating the power plant connected to a control member configured to control the actuator. The control device includes a single control member, an immobilization unit configured to immobilize all the movable parts which are deactivated only when the regulation system controls the positional change of the movable part or parts and a determination unit configured to determine the actual position of the movable part or parts. | 10-17-2013 |
| 20130276426 | Cooling Jacket with Porous Matrix - The fluids and heat transfer theory for regenerative cooling of a rocket combustion chamber with a porous media coolant jacket is presented. This model is useful for calculating temperature distributions in a coolant fluid and combustion chamber or heat source as well as the associated fluid pressure drop through the coolant jacket. This model for fluids and heat transfer theory can be used to design a regeneratively cooled rocket engine. | 10-24-2013 |
| 20130318944 | Jet Exhaust Noise Reduction - Reducing jet noise by weakening Mach cones in a jet exhaust gas streamtube. The Mach cones are weakened by modifying exhaust gas flow in a longitudinal axial core of the exhaust gas streamtube. | 12-05-2013 |
| 20130340406 | FAN STAGGER ANGLE FOR GEARED GAS TURBINE ENGINE - A gas turbine engine includes a spool, a turbine coupled with the spool, a propulsor coupled to be rotated about an axis by the turbine through the spool and a gear assembly coupled between the propulsor and the spool such that rotation of the spool results in rotation of the propulsor at a different speed than the spool. The propulsor includes a hub and a row of propulsor blades that extends from the hub. Each of the propulsor blades has a span between a root at the hub and a tip, and a chord between a leading edge and a trailing edge such that the chord forms a stagger angle α with the axis. The stagger angle α is less than 62° at all positions along the span, with said hub being at 0% of the span and the tip being at 100% of the span. | 12-26-2013 |
| 20130340407 | CLUSTERED, FIXED CANT, THROTTLEABLE ROCKET ASSEMBLY - A clustered, fixed cant, throttleable rocket assembly is used to propel and a steer a vessel in terrestrial or extraterrestrial applications. The fixed cant of each of at least three individual rockets in the cluster provides the steering input to the overall assembly. More specifically, by changing the propellant flow rate to the individual rocket engines relative to one another, the overall thrust vector of the rocket assembly may be selected to provide a desired steering input to the vessel. A measured vessel orientation may be compared with a desired vessel orientation to determine what steering input is required to achieve the desired vessel orientation. | 12-26-2013 |
| 20140026535 | HIGH SPECIFIC IMPULSE SUPERFLUID AND NANOTUBE PROPULSION DEVICE, SYSTEM AND PROPULSION METHOD - A propulsion device including a chamber that stores a superfluid, a substrate coupled to a portion of the chamber, a plurality of orifices extending through the substrate, each of the plurality of orifices having a first end and a second end opposite the first end, the first end disposed in an interior of the chamber and the second end disposed outside the chamber; and a pressure source that generates a pressure differential between the first end of each of the plurality of orifices and the second end of each of the plurality of orifices. | 01-30-2014 |
| 20140026536 | VARIABLE AREA FAN NOZZLE POSITION AND SKEW SENSING - Systems, devices, and methods are presented for detecting a misaligned or otherwise skewed variable area fan nozzle (VAFN) that is mounted on a thrust reverser sleeve of an aircraft engine. Rods project from different parts around the arcuate VAFN, the rods having patterns at their ends. In normal operation when the thrust reverser is stowed, the rods project over the thrust reverser into a fixed area of the engine. Sensors in the fixed area determine the position of the VAFN from the patterns on the rods. When the thrust reverser deploys, the rods are pulled aft with the VAFN, which is mounted on the thrust reverser sleeve, and separate from the sensors. When the thrust reverser stows, the rods move forward with the VAFN and re-engage with the sensors. Repeating patterns on the rods allow for simple, relatively low-cost sensors to read their relative positions. | 01-30-2014 |
| 20140060004 | TILTROTOR VECTORED EXHAUST SYSTEM - The exhaust system is located on each nacelle of a tiltrotor aircraft. The exhaust system includes a vector nozzle that is selectively rotatable in relation to each nacelle in order to achieve certain performance objectives. The vector nozzle can be oriented to provide maximum flight performance, reduce infrared (IR) signature, or even to reduce/prevent ground heating. | 03-06-2014 |
| 20140090358 | REPLACEABLE THRUST GENERATING STRUCTURES ATTACHED TO AN AIR VEHICLE - A replaceable thrust generating structure for an air vehicle, which includes one or more mounts, which are attachable to the air vehicle, and a thrust generating structure attached to each of the one or more mounts, and wherein the thrust generating structure includes an NMSET element. | 04-03-2014 |
| 20140090359 | MIXER THAT PERFORMS RECIPROCATING ROTARY MOTION FOR A CONFLUENT-FLOW NOZZLE OF A TURBINE ENGINE, AND A METHOD OF CONTROLLING IT - The invention provides a mixer for a confluent-flow nozzle of a turbine engine, the mixer comprising an annular cap designed to be centered on a longitudinal axis of the nozzle and having a stationary upstream portion and a downstream portion that is movable in rotation about the longitudinal axis relative to the stationary portion, the movable portion of the cap terminating at its downstream end in inner lobes alternating circumferentially with outer lobes, and means for imparting reciprocating rotary motion to the movable portion of the cap. | 04-03-2014 |
| 20140102074 | APPARATUS AND METHOD FOR A SOLID CATALYST AND FLUID DYNAMIC ERUPTION REACTION - An apparatus and method for use in conducting an eruption reaction are disclosed. The apparatus includes a catalytic solids container with a mouth and fluid egress opening and a trigger device or mechanism that allows for the controlled release of a catalytic solid into an eruptible fluid. The catalytic solids container may be adapted to be coupled to a container for an eruptible fluid. | 04-17-2014 |
| 20140109550 | JET EXHAUST PISTON ENGINE - This invention is a piston in cylinder engine using water injection into a relative vacuum heated to steam by expanding in the cylinder and by an electric arc or other heat source. The resulting steam explosion applies a work force on the piston. The piston has jet nozzles uncovered at the end of its work stroke to jet the piston to help propel it during the return stroke and to form a vacuum in place of the usual compression stroke. The piston has a cover plate with tapered pins depending into jet nozzles through the piston to block the jet nozzles during the main work stroke. | 04-24-2014 |
| 20140123625 | ENGINE CONTROL PARAMETER TRIMMING - A method of re-trimming an engine control parameter. Measure an engine parameter over time. Calculate change in the measured parameter per engine cycle. Determine change of thrust due to the calculated change in the measured parameter from a reference database. Define a new trim value for the engine control parameter that is a function of the change of thrust. Apply the new trim value to the engine control parameter. | 05-08-2014 |
| 20140137538 | Fast Response Bypass Engine - A gas turbine engine has a compressor, a fan for delivering air into the compressor and into a bypass duct, a combustion section and a turbine section. A control for the gas turbine engine is programmed to change a fueling level and position at least one effector that may be moved to unique positions in a coordinated fashion upon receipt of a command to change thrust. The thrust provided by the engine is changed without a reduction in an airflow stability margin compared to a thrust change commanded only by a fueling change. Some aspects of the positioning are transitory. | 05-22-2014 |
| 20140150403 | THRUST REVERSER SYSTEM WITH TRANSLATING-ROTATING CASCADE AND METHOD OF OPERATION - A thrust reverser system and operation suitable for high-bypass turbofan engines. The thrust reverser system includes a cascade system adapted to be translated with a translating cowl in an aft direction of an engine to expose a circumferential opening. The cascade system is deployed from a stowed position as the translating cowl and the cascade system are translated in the aft direction. During deployment of the cascade system, a fore end thereof translates in the aft direction and an aft end thereof initially translates in the aft direction and then subsequently rotates about the fore end so that further translation of the cascade segment in the aft direction causes the cascade segment to move to a deployed position and divert bypass air within a bypass duct of the engine through the circumferential opening. | 06-05-2014 |
| 20140157752 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A method of designing a gas turbine engine includes providing a fan section including a fan; driving the fan section via a gear arrangement; providing a compressor section, including both a first compressor and a second compressor; and driving the compressor section and the gear arrangement via a turbine section. The pressure ratio across the first compressor is greater than or equal to about 7. | 06-12-2014 |
| 20140157753 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A method of designing a gas turbine engine includes providing a fan section including a fan; driving the fan section via a gear arrangement; providing a compressor section, including both a first compressor and a second compressor; and driving the compressor section and the gear arrangement via a turbine section. An overall pressure ratio, being provided by the combination of a pressure ratio across the first compressor and a pressure ratio across the second compressor, is greater than or equal to about 35. The pressure ratio across the first compressor is greater than or equal to about 7. | 06-12-2014 |
| 20140165532 | System and method for improving the efficiency of a jet engine - A method of improving efficiency comprises providing an impulse reaction engine having an exhaust and providing a cascading pipe comprising an exhaust intake having a first diameter, approximately cylindrical walls having a second diameter greater than the first diameter, the approximately cylindrical walls having a first end and a second end, the first end connected to the exhaust intake via a plate comprising at least one orifice, and a fluid injector connected to the at least one orifice and configured to direct fluid in a direction substantially parallel to the approximately cylindrical walls and toward the second end. The method further comprises connecting the exhaust intake of the cascading pipe to the exhaust of the impulse reaction engine and injecting fluid via the fluid injector into the cascading pipe so as to cool exhaust gases emitted from the exhaust of the impulse reaction engine. | 06-19-2014 |
| 20140174054 | FEED SYSTEM AND A METHOD OF SUPPRESSING THE POGO EFFECT - A rocket engine feed system, includes a feed circuit including a device for varying a gas volume in the feed circuit to suppress the POGO effect. A method of suppressing the POGO effect varies at least one hydraulic resonant frequency by varying a rate at which gas is injected into the feed circuit. | 06-26-2014 |
| 20140182265 | Rocket Propulsion Systems, and Related Methods - In an aspect of the invention, a system for rocket propulsion includes a heater operable to generate thermal energy from energy supplied from a non-chemical energy source, and to supply the thermal energy to a non-cryogenic fuel to thermally decompose the fuel into components that include at least a first component and a second component. The rocket propulsion system also includes a combustion chamber and a nozzle. The combustion chamber is operable to receive an oxidizer and at least a portion of the thermally decomposed fuel, and allow the two to combust. The nozzle generates thrust by directing the products of the combustion out of the system. | 07-03-2014 |
| 20140196436 | OPERATION STABILIZATION METHOD AND OPERATION STABILIZATION APPARATUS FOR SUPERSONIC INTAKE - An object of the present invention is to provide a technique of enlarging a stable operating range of an intake in accordance with an operating condition of an engine without using a complicated control system so that a wide operating range of the engine can be covered. In an operation stabilization method for a supersonic intake according to the present invention, an enlarged duct between a cowl and a ramp of the intake is divided by a splitter plate such that an opening angle of the enlarged duct decreases. Further, in the operation stabilization method for a supersonic intake according to the present invention, when the duct is divided by the splitter plate, the splitter plate is disposed such that, from among a cowl side duct and a ramp side duct, cross-section variation in one duct, in which an effect of a flow is larger, is reduced within an allowable range of total pressure loss in the other duct. | 07-17-2014 |
| 20140202132 | HIGH PERFORMANCE LIQUID ROCKET PROPELLANT - Disclosed is a process of fueling a rocket engine or air-breathing engine for a hypersonic vehicle with a high performance hydrocarbon fuel characterized by a hydrogen content greater than 14.3% by weight, a hydrogen to carbon atomic ratio greater than 2.0 and/or a heat of combustion greater than 18.7 KBtu/lb. The disclosed fuels generally have a paraffin content that is at least 90% by mass and a C | 07-24-2014 |
| 20140216002 | GAS TURBINE ENGINE HAVING SLIM-LINE NACELLE - A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a nacelle assembly that includes an inlet lip section and an inlet internal diffuser section downstream of the inlet lip section. A variable area fan nozzle is positioned near an aft segment of the nacelle assembly, the variable area fan nozzle adaptable to move between a first position having a first discharge airflow area and a second position having a second discharge airflow area greater than the first discharge airflow area. At least one boundary layer control device is positioned near one of the inlet lip section and the inlet internal diffuser section. A controller is configured to move the variable area fan nozzle from the first position to the second position and to actuate the at least one boundary layer control device to introduce an airflow in response to an operability condition. | 08-07-2014 |
| 20140250861 | AIRCRAFT POWER PLANT - An aircraft power plant is disclosed having a plurality of bladed rotors in flow communication driven by separate work producing devices. The work producing devices can take a variety of forms including an internal combustion engine and electric motor, for example. The bladed rotors can be associated with an aircraft pylon and can be driven independently to separate operating conditions to provide optimum performance. For example, the bladed rotors can be driven to separate operating conditions that improve a noise signature or performance of the aircraft. | 09-11-2014 |
| 20140260180 | Gas Turbine Engine with Stream Diverter - In accordance with one aspect of the disclosure, a stream diverter for a gas turbine engine is disclosed. The stream diverter may include a first air duct, a second air duct, a third air duct, and a door operatively associated with the second and third air ducts of the gas turbine engine. The door may have at least an open position allowing air from the second air duct to flow into the third air duct and a closed position preventing air from flowing between the ducts. | 09-18-2014 |
| 20140260181 | JET PROPULSION DEVICE AND FUEL SUPPLY METHOD - A reaction propulsion device in which a first feed circuit for feeding a main thruster with a first propellant includes a branch connection downstream from a pump of a first turbopump, which branch connection passes through a first regenerative heat exchanger and a turbine of a first turbopump, and in which a second feed circuit for feeding the main thruster with a second propellant includes, downstream from a pump of a second turbopump, a branch-off passing through a second regenerative heat exchanger and a turbine of the second turbopump. At least one secondary thruster is connected downstream from the turbines of the first and second turbopumps. | 09-18-2014 |
| 20140283499 | DEVICE AND A METHOD FOR FEEDING A ROCKET ENGINE PROPULSION CHAMBER - The invention relates to a device and a method for feeding a propulsion chamber | 09-25-2014 |
| 20140290211 | TURBINE ENGINE INCLUDING BALANCED LOW PRESSURE STAGE COUNT - A turbine engine includes at least a compressor section and a turbine section, each having at least a first and second portion. A ratio of turbine section second portion stages to compressor section second portion stages is less than or equal to 1. | 10-02-2014 |
| 20140305097 | FAN VARIABLE AREA NOZZLE WITH CABLE ACTUATOR SYSTEM - An assembly for pivoting a flap according to an exemplary aspect of the present disclosure includes, among other things, a structure is mounted at least partially around an axis. The structure is attached to a pivotable flap arranged to define a nozzle area. A cable passes through an orifice defined by the flap. An actuator system is operable to mechanically retract the cable therein to lessen the nozzle area and mechanically extend the cable to enable the flow to increase the nozzle area. The actuator system is engaged with the cable. A segment of the cable, opposite the actuator system, is attached to a fixed structure. A method of providing a variable fan exit area is also disclosed. | 10-16-2014 |
| 20140311121 | PULSE DETONATION ENGINE HAVING A SCROLL EJECTOR ATTENUATOR - The engine ( | 10-23-2014 |
| 20140325956 | SYSTEM, APPARATUS, AND METHOD FOR THRUST VECTORING - In various embodiments, the thrust vectoring systems described herein create variable reverse thrust during a landing event. The reverse thrust may be varied based on manual inputs, dynamically changing operating events, or a landing duty cycle. In various embodiments, the thrust vectoring systems comprise a movable shelf that is capable of adjusting a directing a fluid flow to create a variable reverse thrust which may reduce the risk of foreign object ingestion in the engine at lower ground speeds. | 11-06-2014 |
| 20140345252 | SYSTEM, APPARATUS, AND METHOD FOR A VIRTUAL BLOCKER - A virtual blocker door for use in an aircraft thrust reverser is provided. The virtual blocker door may be capable of creating a wall of air to inhibit fan air flow through a fan air duct. The wall of air may direct the fan air flow through a cascade to provide a reverse thrust. Moreover, the virtual blocker door provides for improved aerodynamic efficiency in the fan air duct by replacing traditional brackets, drag links, and doors with a wall of air. | 11-27-2014 |
| 20140360157 | ROCKET VEHICLE WITH INTEGRATED ATTITUDE CONTROL AND THRUST VECTORING - A rocket vehicle includes a controller that integrates operation of a variable-vector main thruster and attitude control thrusters. When the main thruster is firing and roll is commanded, the controller can provide roll moment by firing only a single attitude control thruster, while changing the thrust vector of the main thruster to offset any pitch/yaw moments induced by the firing of the single attitude control thruster. The single attitude control thruster may be a thruster on the leeward side of the rocket vehicle. Since there is a lower wall pressure on the leeward side of the rocket vehicle, the thruster efficiency is improved by accomplishing roll by use of a single thruster (which may be one of a pair of thrusters used to achieve roll in one direction). A significant reduction in fuel use may be accomplished. | 12-11-2014 |
| 20150007549 | Rotary Turbo Rocket - A turbojet is combined with a co-axially integrated rotary rocket to form a propulsion system called a Rotary Turbo Rocket that can function as a turbojet, as an afterburning turbojet, as an Air Turbo Rocket, or as a rotary rocket. The Rotary Turbo Rocket can operate in any of these propulsion modes singularly, or in any combination of these propulsion modes, and can transition continuously or abruptly between operating modes. The Rotary Turbo Rocket can span the zero to orbital flight velocity speed range and/or operate continuously as it transitions from atmospheric to space flight by transitioning between operating modes. | 01-08-2015 |
| 20150007550 | Combined Cycle Integrated Combustor and Nozzle System - An engine that operates and produces the entire required vehicle thrust below Mach 4 is useful for a Hypersonic combined cycle vehicle by saving vehicle and engine development costs. One such engine is a combined cycle engine having both a booster and a dual mode ramjet (DMRJ). The booster and the DMRJ are integrated to provide effective thrust from Mach 0 to in excess of Mach 4. As the booster accelerates the vehicle from Mach 0 to in excess of Mach 4, from Mach 0 to about Mach 2 incoming air delivered to the DMRJ is accelerated by primary ejector thrusters that may receive oxidizer from either on-board oxidizer tanks or from turbine compressor discharge air. As the TBCC further accelerates the vehicle from about Mach 0 to in excess of Mach 4 exhaust from the turbine and exhaust from the DMRJ are combined in a common nozzle disposed downstream of a combustor portion of the DMRJ functioning as an aerodynamic choke. | 01-08-2015 |
| 20150013301 | TURBINE ENGINE INCLUDING BALANCED LOW PRESSURE STAGE COUNT - A turbine engine includes at least a compressor section and a turbine section, each having at least a first and second portion. A ratio of turbine section second portion stages to compressor section second portion stages is less than or equal to 1. | 01-15-2015 |
| 20150013302 | ENGINE PROPULSION SYSTEM - An engine propulsion system is configured to utilize bursts of media in order to create mechanical energy. The engine propulsion system includes at least one cannon, wherein each cannon is configured to displace the media and further includes a firing pin casing configured to accommodate a firing pin. The firing pin is configured to create the mechanical energy when moved thus allowing the media to exit the cannon. | 01-15-2015 |
| 20150013303 | ENGINE PROPULSION SYSTEM - An engine propulsion system is configured to utilize bursts of media in order to create mechanical energy. The engine propulsion system includes at least one cannon, wherein each cannon is configured to displace the media and further includes a firing pin casing configured to accommodate a firing pin. The firing pin is configured to create the mechanical energy when moved thus allowing the media to exit the cannon. | 01-15-2015 |
| 20150013304 | ACOUSTICALLY TRIGGERED NANO/MICRO-SCALE PROPULSION DEVICES - Techniques, devices and systems are disclosed for implementing acoustically triggered propulsion of nano- and micro-scale structures. In one aspect, an ultrasound responsive propulsion device includes a tube that includes one or more layers including an inner layer having an electrostatic surface, and an ultrasound-responsive substance coupled to the inner layer and configured to form gaseous bubbles in a fluid in response to an ultrasound pulse, in which the bubbles exit the tube to propel the tube to move in the fluid. | 01-15-2015 |
| 20150047315 | GAS TURBINE ENGINE FLOW DUCT HAVING INTEGRATED HEAT EXCHANGER - A gas turbine engine flow duct comprising a flow duct disposed along an engine centerline of the gas turbine engine and defining a stream flow passage, and first and second rows of heat exchangers disposed along the engine centerline of the gas turbine engine and integrated in the flow duct in fluid communication with the stream flow passage of the flow duct. | 02-19-2015 |
| 20150052874 | MICRO-CATHODE THRUSTER AND A METHOD OF INCREASING THRUST OUTPUT FOR A MICRO-CATHODE THRUSTER - A magnetically enhanced micro-cathode thruster assembly provides long-lasting thrust. The micro-cathode thruster assembly includes a tubular housing, a tubular cathode, an insulator, an anode and a magnetic field. The tubular housing includes an open distal end. The tubular cathode is housed within the housing and includes a distal end positioned proximate the open distal end of the housing. The insulator is in contact with the cathode forming an external cathode-insulator interface. The anode is housed within the housing, proximate the open distal end of the housing. The magnetic field is positioned at or about the external cathode-insulator interface and has magnetic field lines with an incidence angle of about 20 to about 30 degrees and preferably about 30 degrees relative to the external cathode-insulator interface. | 02-26-2015 |
| 20150101308 | ENGINE - The present disclosure relates to an engine having two modes of operation—air breathing and rocket—that may be used in aerospace applications such as in an aircraft, flying machine, or aerospace vehicle. The engine's efficiency can be maximized by using a precooler arrangement to cool intake air in air breathing mode using cold fuel used for the rocket mode. By introducing the precooler and certain other engine cycle components, and arranging and operating them as described, problems such as those associated with higher fuel and weight requirements and frost formation can be alleviated. | 04-16-2015 |
| 20150113941 | TRANSLATING OUTER COWL FLOW MODULATION DEVICE AND METHOD - A flow control device includes a first axially extending flow control surface, a second axially extending flow control surface radially offset from the first surface to define a gas flow path therebetween, the gas flow path having a downstream flow path exit, and a third axially extending flow control surface radially offset from the first surface and capable of axially translating with respect to the first and second surfaces for modifying the gas flow path and selectively closing the flow path exit. A turbofan engine includes a core flow passage, a fan bypass passage located radially outward from the core flow passage, a third stream bypass passage located radially outward from the fan bypass passage, and a flow control device that dynamically regulates the third stream bypass passage, allowing fluid flowing through the third stream bypass passage to provide thrust to the turbofan engine and reduce afterbody drag. | 04-30-2015 |
| 20150121839 | TANDEM THRUST REVERSER WITH MULTI-BAR LINKAGE - One embodiment includes a pivot thrust reverser. The pivot thrust reverser includes a first tandem pivot door subassembly comprising an inner panel and an outer panel. The inner panel and the outer panel are connected so as to rotate simultaneously about respective pivot axises that are each positionally fixed axises relative to the gas turbine engine assembly. A second tandem pivot door subassembly is included, spaced from the first tandem pivot door subassembly and comprising an inner panel and an outer panel. The inner panel and the outer panel are connected so as to rotate simultaneously about respective pivot axises that are each positionally fixed axises relative to the gas turbine engine assembly. | 05-07-2015 |
| 20150121840 | AIRCRAFT NOZZLE SYSTEM - A nozzle is provided that is capable providing flowpaths for a combined cycle aircraft propulsion system that in one form includes a gas turbine engine and a ramjet. The gas turbine engine produces an exhaust flow that is offset from an exhaust flow from the ramjet. The two streams can be flowed independent of each other or together depending on the application and relevant portion of a flight envelope. The nozzle includes a movable portion that can selectively open and close an exhaust flowpath for the gas turbine engine. The nozzle includes a surface that provides expansion for both low speed (gas turbine engine) flow and high speed (ramjet) flow. | 05-07-2015 |
| 20150121841 | FLOW CONTROL DEVICE FOR A THREE STREAM TURBOFAN ENGINE - A gas turbine engine includes a core engine, a first bypass passage disposed about the core engine and a second bypass passage disposed about the first bypass passage. A flow control is disposed within the second bypass for controlling bypass airflow through the second bypass. The flow control translates axially between an open position and a closed position to vary and control airflow through the second bypass passage. | 05-07-2015 |
| 20150121842 | PNUEMATIC SYSTEM FOR AN AIRCRAFT - A pneumatic system | 05-07-2015 |
| 20150128561 | THREE STREAM, VARIABLE AREA, VECTORABLE NOZZLE - An exhaust nozzle for a gas turbine engine may include a plurality of flap trains in the exhaust stream of the gas turbine engine. The flap trains are operable to selectively control three separate flow paths of gas that traverse the engine. A first stream of is the core airflow. The second stream of air is peeled off of the first stream to form a low pressure fan bypass air stream. The third stream of air traverses along the engine casing and is passed over a flap assembly to aid in cooling. The flaps are operable converge/diverge to control the multiple streams of air. | 05-14-2015 |
| 20150135678 | GAS TURBINE ENGINE WITH NOISE ATTENUATING VARIABLE AREA FAN NOZZLE - A nacelle assembly for a high-bypass gas turbine engine includes a core nacelle defined about an engine centerline axis. A fan nacelle is mounted at least partially around the core nacelle to define a fan bypass flow path. A variable area fan nozzle is in communication with the fan bypass flow path. The variable area fan nozzle has a first fan nacelle section and a second fan nacelle section. The second fan nacelle section is axially movable relative to the first fan nacelle section to define an auxiliary port at a non-closed position to vary a fan nozzle exit area and adjust fan bypass airflow. The second fan nacelle section includes an acoustic system that has an acoustic impedance located on a radially outer surface. | 05-21-2015 |
| 20150300292 | SYSTEM OF SUPPORT THRUST FROM WASTED EXHAUST - Systems and methods for influencing thrust of a gas turbine engine are disclosed. A system includes a plurality of pipes positioned at an exhaust nozzle of the gas turbine engine. Each one of the plurality of pipes includes: an inlet inside the exhaust nozzle; an outlet inside the exhaust nozzle; and an intermediate portion outside the exhaust nozzle. | 10-22-2015 |
| 20150308339 | GAS TURBINE ENGINE WITH LOWER BIFURCATION HEAT EXCHANGER - A gas turbine engine includes a nacelle, a lower bifurcation structure, and a heat exchanger. The nacelle extends circumferentially around an engine core and defines a fan bypass duct that is substantially annular between an inner wall and an outer wall. The lower bifurcation structure extends between the inner wall and the outer wall, bifurcating the fan bypass duct. The lower bifurcation structure defines a bifurcation duct extending along a central axis of the lower bifurcation structure. A heat exchanger is positioned in the bifurcation duct. | 10-29-2015 |
| 20150354505 | PROPELLANT FEED CIRCUIT AND A COOLING METHOD - The invention relates to the aerospace field, and in particular to the field of vehicles propelled by rocket engines. In particular, the invention relates to a feed circuit ( | 12-10-2015 |
| 20160001897 | CONTROLLING A PROPELLANT DISTRIBUTION IN A SPACECRAFT PROPELLANT TANK - A system for controlling a distribution of propellant in a propellant tank assembly for a spacecraft comprises a body for containing the propellant, a plurality of thermal tomography elements, including a plurality of temperature-control elements and a plurality of temperature sensors, disposed around the body for detecting the distribution of the propellant inside the body; and a tomography element control module arranged to control the plurality of temperature-control elements to redistribute the propellant inside the propellant tank body by heating and/or cooling the propellant. In an embodiment, the propellant tank body includes a propellant management device inside the body and the tomography elements are disposed in proximity to the propellant management device. Tomography data can be obtained from the plurality of tomography elements, and a distribution of propellant within the propellant tank body can be determined based on the obtained tomography data. | 01-07-2016 |
| 20160001899 | Gas Envelope Propulsion System and Related Methods - Disclosed are exemplary embodiments of systems and methods for gas envelope propulsion. In an exemplary embodiment, a propulsion system generally includes a power plant, and an inelastic gas envelope configured to receive a gas independently of the power plant. The power plant is configured to provide thrust for moving the gas envelope skyward, and the gas envelope is further configured to expel at least some of the gas to provide additional thrust. | 01-07-2016 |
| 20160053718 | Thrust Reverse Variable Area Fan Nozzle - A thrust reverse variable area nozzle system for a nacelle of an aircraft engine system may include a reverse thrust opening disposed in the nacelle, and a thrust reverser door pivotally movable relative to the nacelle for selectively covering the reverse thrust opening, wherein the thrust reverser door is pivotally movable between a first position for completely covering the reverse thrust opening, a second position for partially uncovering a forward portion of the reverse thrust opening and discharging a bypass airflow through the forward portion of the reverse thrust opening in a forward direction, and a third position for partially uncovering an aft portion of the reverse thrust opening and discharging the bypass airflow through the aft portion of the reverse thrust opening in an aft direction. | 02-25-2016 |
| 20160084199 | JET EXHAUST NOISE REDUCTION - Reducing jet noise by weakening Mach cones in a jet exhaust gas streamtube. The Mach cones are weakened by modifying exhaust gas flow in a longitudinal axial core of the exhaust gas streamtube. | 03-24-2016 |
| 20160097346 | ELECTRIC BOOST ACTUATION SYSTEM FOR TRANSLATING RINGS - An actuation system for a gas turbine engine oriented about a centerline includes a translating ring, at least one hydraulic actuator, and at least one electric actuator. The translating ring is oriented about the centerline and configured to move axially along the centerline. The at least one hydraulic actuator is configured to provide a first mechanical force to move the translating ring along the centerline. The at least one electric actuator is configured to provide a second mechanical force to move the translating ring in the axial direction. The at least one electric actuator is controlled to provide the second mechanical force upon determination of an operating condition. | 04-07-2016 |
| 20160108851 | THRUST VECTORING APPARATUS, THRUST VECTORING METHOD, AND FLYING BODY - A first jet tab and a second jet tab are symmetrically arranged with respect to a first symmetry plane and have a symmetrical shape with respect to the first symmetry plane, and are symmetrically driven with respect to the first symmetry plane by a driving section. A distance between a tip sensation of the first jet tab and a first rotation axis is larger than a distance between the first rotation axis and the first symmetry plane. A distance between a tip section of the second jet tab and a second rotation axis is larger than a distance between the second rotation axis and the first symmetry plane. | 04-21-2016 |
| 20160108855 | DUAL MODE CHEMICAL ROCKET ENGINE, AND DUAL MODE PROPULSION SYSTEM COMPRISING THE ROCKET ENGINE - The invention relates generally to dual mode bipropellant chemical rocket propulsion systems to be used in aerospace applications for 1) orbit raising, orbit manoeuvres and maintenance, attitude control and deorbiting of spacecraft, and/or 2) propellant settling, attitude and roll control of missiles, launchers and space planes. The present invention also relates to a dual mode chemical rocket engine for use in such systems. The engine uses low-hazardous storable liquid propellants and can be operated either in monopropellant mode or in bipropellant mode. The monopropellants used are a low-hazard liquid fuel-rich monopropellant, and hydrogen peroxide, respectively. | 04-21-2016 |
| 20160115905 | TWO-PULSE GAS GENERATOR AND OPERATION METHOD THEREOF - A gas generator includes an outer propellant, an inner propellant arranged inside the outer propellant, and a barrier membrane which isolates the outer propellant from the inner propellant. A forward end surface of the inner propellant faces a combustion space. A side surface of the inner propellant is isolated from the combustion space. | 04-28-2016 |
| 20160115906 | DUAL MODE CHEMICAL ROCKET ENGINE AND DUAL MODE PROPULSION SYSTEM COMPRISING THE ROCKET ENGINE - The invention relates generally to dual mode bipropellant chemical rocket propulsion systems to be used in aerospace applications for 1) orbit raising, orbit manoeuvres and maintenance, attitude control and deorbiting of spacecraft, and/or 2) propellant settling, attitude and roll control of missiles, launchers and space planes. The present invention also relates to a dual mode chemical rocket engine for use in such systems. The engine uses low-hazardous storable liquid propellants and can be operated either in monopropellant mode or in bipropellant mode. The monopropellants used are a low-hazard liquid fuel-rich monopropellant, and a low-hazard liquid oxidizer-rich monopropellant, respectively. | 04-28-2016 |
| 20160122027 | AIRCRAFT SYSTEM - An aircraft capable of operating at a variety of speeds includes a power plant and an auxiliary turbine. The auxiliary turbine can be a ram air turbine used to expand and cool an airflow and provide work. The cooled airflow from the auxiliary turbine can be used in a heat exchange device such as, but not limited to, a fuel/air heat exchanger. In one embodiment the cooled airflow can be used to exchange heat with a compressor airflow being routed to cool a turbine. Work developed from the auxiliary turbine can be used to power a heating device and rotate a device to add work to a shaft of the aircraft power plant. In one form the aircraft power plant is a gas turbine engine and the work developed from the auxiliary turbine is used to heat a combustor flow or to drive a shaft that couples a turbine and a compressor. | 05-05-2016 |
| 20160123180 | OVER SPEED MONITORING USING A FAN DRIVE GEAR SYSTEM - A control system for turbofan engine includes a first sensor measuring rotation of a first shaft at a first location and a fan shaft sensor measuring a speed of a fan shaft. A controller utilizes measurements of a first speed of the first shaft from the first sensor and a second speed of the fan shaft driven by a geared architecture and rotating at a speed different than the first shaft. The controller determines that one of the first shaft and the fan shaft are outside predetermined deformation limits responsive to a difference between an actual difference between the first and second speeds and a calculated expected difference between speeds of the first shaft and the fan shaft. | 05-05-2016 |
| 20160131085 | STORED PRESSURE DRIVEN CYCLE - A propulsion system according to an exemplary aspect of the present disclosure includes, among other things, a pressurant selectively released from a pressure tank to drive a pump to sustain propellant flow for main combustion. | 05-12-2016 |
| 20160138523 | Optimal Thrust Control of an Aircraft Engine - 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 nozzle scheduler for determining an exhaust nozzle position goal based on a nozzle schedule of exhaust nozzle positions related to flight conditions. The control system may further include a control module for determining a control command for the gas turbine engine. The control command may include, at least, a fuel flow command and an exhaust nozzle position command and the control command may be based on, at least, the exhaust nozzle position goal and an estimated thrust value. The control system may further include an actuator for controlling the gas turbine engine based on the control command. | 05-19-2016 |
| 20160138524 | Monopropellant Driven Hydraulic Pressure Supply - A liquid propellant driven hydraulic pressure supply device may include an elongated body having an internal bore extending from a power end to a discharge end having a discharge port, a hydraulic fluid disposed in the bore between a piston and the discharge end and a liquid propellant gas generator connected to the power end. | 05-19-2016 |
| 20160146107 | Variable Supersonic Engine Inlet - Systems and methods for generating an oblique shock in a supersonic inlet are disclosed. The system can comprise an inlet with a slot disposed at an oblique angle to the main incoming air stream. High-pressure air can be provided through the slot into the main air stream. The high-pressure air can be introduced at a high enough pressure ratio—i.e., the ratio of pressure of the air stream from the slot to the pressure for the main flow—such that an aerodynamic ramp is created in the main air flow. The aerodynamic ramp, in turn, can cause one or more oblique shock waves to eventually slow the main air stream velocity to a subsonic speed prior to the face of the engine. Systems and methods for controlling the slot pressure ratio to create these shocks are also disclosed. | 05-26-2016 |
| 20170234233 | BOWED ROTOR START MITIGATION IN A GAS TURBINE ENGINE USING AIRCRAFT-DERIVED PARAMETERS | 08-17-2017 |
| 20180023512 | Rocket Fueling Systems and Methods | 01-25-2018 |
| 20190145348 | SOUND-ATTENUATING HEAT EXCHANGERS AND METHODS OF UTILIZING THE SAME | 05-16-2019 |
