264 results about "Flight envelope" patented technology

An air vehicle incorporating a hybrid propulsion system. The system includes a gas turbine engine as a first motive power source, and one or more battery packs as a second motive power source. Through selective coupling to a DC electric motor that can in turn be connected to a bladed rotor or other lift-producing device, the motive sources provide differing ways in which an aircraft can operate. In one example, the gas turbine engine can provide operation for a majority of the flight envelope of the aircraft, while the battery packs can provide operation during such times when gas turbine-based motive power is unavailable or particularly disadvantageous. In another example, both sources of motive power may be decoupled from the bladed rotor such that the vehicle can operate as an autogyro.

A system and method for automatically varying the flight envelope of an aircraft based upon the material health of the aircraft and the flight environment is provided. The system includes a plurality of structural health monitoring and load sensors that determine the approximate size and the approximate location of the damage. The system performs residual strength calculations for individual aircraft components to determine the overall aircraft residual strength. The system uses these calculations to determine a maximum flight envelope based on the overall aircraft residual strength, and transmits this information to the flight controller and optionally to the pilot.

A Blended Wing Body SUAV and MUAV is disclosed having a novel airfoil profile, wing configuration, rigging and tractor pull propeller placement that provide improved stability and safety characteristics over prior art SUAVs and MUAVs of comparable size and weight. This unique blended wing design includes wing twist on the outboard wing and an inverted “W” shaped planform to provide lateral and longitudinal stability, and smooth, even flight characteristics throughout the range of the expected flight envelope. These flight characteristics are crucial to providing a stable reconnaissance platform with favorable stall speeds, an increased payload and the ability to hand launch without the danger of exposing ones hands or wrist to a propeller.

Smart instrumentation and monitoring provide primary engine instrumentation for an aircraft. The instrumentation generates real time engine power available data for the entire power regime and flight envelope of the aircraft while engine monitoring functions track, record and output performance usage data, trend data, and flight profile data. Generated data is provided to an aircraft flight manual and mission management module that automatically calculates and generates aircraft operational and performance data needed to safely operate the aircraft. The generated data is selectively displayed on flat-panel cockpit displays and/or electronic kneeboard displays in a graphical format that enhances aircrew situational awareness and permits aircrew members to operate the aircraft with its intended margin of safety.

An unmanned air vehicle comprises a fuselage that defines aerodynamic flight surfaces, an engine mounted to the fuselage having an engine shaft arranged to rotate about a longitudinal axis with respect to the fuselage, and a propeller mounted to the engine shaft so as to rotate to thereby provide thrust. The aircraft also comprises a gyroscopic stabilization member coupled to the shaft such that rotation of the engine shaft results in rotation of the gyroscopic member. Thus, there is more stability during the entire flight envelope. In one embodiment, the gyroscopic stabilization member is comprised of a ring that is attached to the outer ends of the blades of the propeller and the ring is also selected so as to have a mass that will result in the gyroscopic stabilization member having a sufficient angular momentum so as to gyroscopically stabilize the aircraft.

The invention discloses turbine-based combined cycle engine super-combustion chamber. Key parts of the super-combustion chamber comprise a shunt ring, a rear area-variable deflation valve, a flow guide tongue piece, a square lobe mixer, a first area fuel ejector, a center cone, a second area fuel injection ring, an evaporation tube flame stabilizer, a heat-isolating screen, a super-combustion chamber cylinder, an outer extending groove-width-variable radial flame stabilizer and an inner extending radial flame stabilizer. An internal flow field of the super-combustion chamber is reasonably distributed, the phenomenon of backflow from a punching duct to a turbofan duct can not occur, flow separation can not occur in the square lobe mixer, and the evaporation tube flame stabilizer can achieve soft ignition of the super-combustion chamber. The whole super-combustion chamber is high in mixing efficiency, little in cold flow resistance, high in combustion efficiency, reasonable in temperature distribution, and capable of meeting use requirements in a whole flight envelope of the super-combustion chamber in each work state and the requirement for boost of an turbofan ramjet.

The invention pertains to the field of 'engine test equipment and test technologies', and relates to an airplane engine dynamic simulation test bed. The airplane engine dynamic simulation test bed consists of a 50-stitch interface definition, a left and right engine full privilege digital electronic controller, an engine electric control unit, an engine vibration detecting instrument, a testing part and a measuring and controlling system. The measuring and controlling system comprises a simulation/dispersion input interface card, measuring and controlling software of an embedded engine simulation model, a sensor simulation unit and an output interface card. The test bad has the advantages of complete function, high intelligent property, fast response time and capability of precisely simulating the dynamic and steady work process and fault monitoring result of the whole flight envelope internal power device system of an airplane by the engine simulation model, an electric simulation technology and a dynamic simulation technology under the condition of no real engine existing. The test bed can be used in the test study of a power device comprehensive control system of the whole system of the airplane and the function study of the airborne testing part.

An electric turbine compressor fan for hybrid propulsion wherein the compressor contains one or more rotor stages (compressor and diffuser), each being driven by one or more electric ring motors, such that the compressor rotor stages are designed and tuned more precisely to the compression ratio to be attained within the turbine design operating characeristics, thrust requirements and flight envelope.

A gun pod, mounted on a fixed-wing aerial vehicle, stores, delivers, controls and supports a controllable movement gun unit. The gun pod includes a flexible gun mount, gun movement actuators, gun movement controllers, a standalone range finder, a standalone processor, and standalone sensors for capturing dynamically environmental data and for controlling the movement of the gun unit. The gun is provided with allowable ranges of movement in the elevation and the azimuth where the ranges are determined in accordance with the flight envelope of the aerial vehicle, the characteristics of the gun unit and the mounting location of the gun pod. The movement of the gun is either controlled manually or automatically.

Supersonic Magnetic Advanced Generation Jet Electric Turbine (S-MAGJET) described herein, and a subsonic derivative, MAGJET, integrate a gas power turbine, superconducting electric power and propulsion generation, and magnetic power flux field systems along with an ion plasma annular injection combustor which utilizes alternative petroleum-based fuel and combustion cycles to create a hybrid turbine turbomachine for aerospace propulsion. The propulsion unit is able to achieve a dramatic increase in horsepower, combustion and propulsion efficiency, and weight reduction. In addition, the turbomachinery structures may be disposed within an exo-skeleton architecture that achieves an increase in thrust to weight ratio with a concomitant increase in fuel efficiency and power generation over traditional gas turbine technology today. The engine continuously adjusts the temperature, pressure and mass airflow requirements using an electromagnetic power management system architecture. Engine performance may be controlled across the entire desired flight envelope, whether subsonic, transonic or supersonic flight conditions. With bypass fan(s), compressor stages and turbine segments controlled electrically in a shaftless turbine design, the S-MAGJET maximizes the propulsion efficiencies over a broader range of operating conditions compared to current art of turbine technology and at much higher thermodynamic and aerodynamic efficiencies.

The invention discloses a synthetic modeling method of rotor and airframe coupling dynamics modality, and belongs to the dynamic design technology of helicopters. The synthetic modeling method of the rotor and airframe coupling dynamics modality is suitable for the dynamic stability and dynamic response analysis modeling of single-rotor and multi-rotor coupling. The synthetic modeling method analyzes the air resonance instability of a rotor blade by aiming at the rotor blade operation pneumatic environment of a flight state stipulated by a flight envelope rule which needs to considers to cover and properly exceed the performance requirements of the helicopter, researches the influence of high-speed high-overload flight state parameters, looks for the structural design parameters of rotors, blades and a steering control system which cause or affect the generation of the air resonance, and determines the instability boundary of the air resonance. Through the sensitivity analysis of the air resonance by the structural design parameters, a structural design parameter range which affects the air resonance is determined, and a design measure for expanding the security boundary of the air resonance is put forward.

Disclosed is a high-power electric motor and its fabrication technology. The motor and its distributed power electronics are all being fully integrated in a conventional turbofan engine. The rotor drives directly (with no gears) the LP shaft of the jet engine while requiring minimal modification to a basic jet engine and without distortion to the nacelle geometry. In principle such a configuration should be suitable for a power level of 10 to 50 MW, which makes it fully capable of providing a standard flight envelope by only using electric energy.

The invention provides a differential adjustable unilateral expansion nozzle. In the invention, a nozzle throat area is bilaterally adjustable and differential adjustment can be carried out to ensure that an engine flow variation requirement is satisfied in a whole flight envelope; an nozzle exit area is adjustable to ensure that a gas ideal complete expansion requirement can be satisfied in a certain flight Mach number range; and the upper web of a nozzle is divided into an adjustable upper web and a fixed upper web between which a certain gap exists and the after body of an aircraft can be directly utilized by the fixed upper web so that that the operating force of a movement mechanism can be reduced. A boundary layer discharged from a front fuselage can be injected into the fixed upperweb through the gap under the flight conditions of transonic speed and low supersonic speed; and therefore, the wall static pressure of the fixed upper web is increased, the thrust loss caused by airflow overexpansion is decreased, meanwhile, the pitching moment of the aircraft is reduced and the nozzle is ensured to always achieve a higher thrust coefficient in the broad flight envelope, in addition, the cooling of the fixed upper web can be strengthened and the thermal load and the infrared signal characteristic of the nozzle are reduced.

The invention discloses an aircraft engine gas path fault fusion diagnosis method. The aircraft engine gas path fault fusion diagnosis method includes the steps that an engine state variable model is built firstly, sensor bias values (NL, NH, P3 and T6) obtained when an engine with a gas path fault is in a quasi-stable state in a flight envelope, then an optimal particle is obtained through a Kalman filter according to an electromagnetism-like mechanism method, and the sensor bias values (NL, NH, P3 and T6) of the engine with the gas path fault and gas path component health parameters are obtained finally. The aircraft engine gas path fault fusion diagnosis method breaks through limitation that a traditional gas path fault diagnosis method can only achieve single fault mode diagnosis, and can achieve health diagnosis of three broad categories of fault modes including the engine sensor fault, the gas path component fault and the sensor and gas path component simultaneous fault.

The invention belongs to the field of environment simulation and control, and relates to a dynamic aviation thermal power testing system and a rapid temperature and pressure control method which are mainly used for the airplane environment control system ground simulation test and are capable of achieving rapid temperature, pressure and flow adjustment and carrying out rapid temperature increasing and reducing and pressure increasing and reducing of an environment control system and the full-flight envelope dynamic simulation test. According to the dynamic aviation thermal power testing system, the project achieving difficulty of a heater is lowered through an indirect heating mode, and the possible problem of dried burning due to direct heating can be avoided. The temperature and pressure are independently controlled, and the temperature and pressure coupling effect is eliminated. A rapid adjusting mechanism is arranged in front of a heat exchanger, the project achieving difficulty of the adjusting mechanism is lowered, accordingly, the reliability of the system is improved, and the manufacturing cost of the system is reduced. The system is further provided with a heat regenerator, part of waste heat can be recycled, the system efficiency is improved, and the consumed energy of the system is reduced.

The invention discloses an unmanned aerial vehicle obstacle avoidance method, which comprises the steps of planning an initial flight path for an unmanned aerial vehicle; detecting position and speed information of an obstacle within a certain distance of the unmanned aerial vehicle by using a sensor in the flying process of the unmanned aerial vehicle; calculating whether collision occurs or not in flight according to the initial path by adopting a collision detection algorithm; working out a collision avoidance path by using a policy iteration algorithm according to a collision region information predicted by the collision detection algorithm and a destination of the unmanned aerial vehicle; correcting the collision avoidance path by using a backtracking method, and preventing the unmanned aerial vehicle from being caught in a local obstacle trap; and performing differential flattening processing on the corrected collision avoidance path according to dynamic performance characteristics of the unmanned aerial vehicle. According to the method provided by the invention, flight envelope restriction requirements of the unmanned aerial vehicle in reality are met, a common problem of local obstacle traps in the path planning algorithm is effectively avoided, and multiple obstacles can be avoided in real time under the premise of meeting the dynamic performance.

The invention relates to a method for determining the speed of an aircraft that is subject to a time constraint. The invention consists no longer in calculating a single CAS/MACH pair during climb/descent but in adapting the speed in a continuous manner to the bounds of curves of minimum V_min and maximum V_max speed defining a flight envelope of the aircraft. The calculation of these speeds is carried out on the basis of constant maximum and minimum speed setpoints and of a coefficient taking into account a deviation to the time constraint.

The invention discloses a dual-redundancy compensation type empennage control plane fault state control method. The method comprises the steps of building a fault state physical model; building a fault state parameterization universal model; calculating a control effect and a hinge moment limiting boundary; determining a residual maneuvering ability and a safe controllable flight envelope; binding an emergency disposal envelope; and performing online use. According to the method, the fault state analysis range is optimized and reduced, so that a maximum safe controllable flight envelope is obtained with a minimum calculation amount; a mathematic model is simplified; a floating yaw angle in a flight process is obtained by adopting a reduced order interpolation method in a control model, so that accurate calculation of the floating-state yaw angle is realized; and the emergency disposal envelope is bound in advance, so that a safe flight range is preset for online emergency. In addition, the influences of the hinge moment of the control plane and the efficiency on the residual maneuvering ability of an aircraft are comprehensively considered, and the height and speed ranges meeting safe flight requirements are preset for fault emergency disposal of the control plane, so that the survivability of the aircraft is improved.