2772 results about "Rocket engine" patented technology

A system for oxygen, hydrogen and carbon mass regeneration and recycling for breathing, and fuel / energy generation purposes, especially for fuel cells and rocket motors, by combination and integration of a photoelectrolytically powered electrochemical and gas handling system with one or more fuel cells.

Method and Apparatus for efficiently applying coating systems to a surface that can survive the thermal gradient that is encountered in high temperature, high heat flux environments such as a rocket engine or like using vapor or cluster deposition process such as a directed vapor deposition (DVD) approach. Method and Apparatus provides electron or other energetic beam technique to evaporate and deposit compositionally and morphologically controlled bond coats at high rate while providing a highly strain tolerant thermal barrier coating that has an improved porosity morphology between columnar grains.

The purpose of the invention is to provide a high-precision liquid or gas rocket engine thrust measuring rack. The thrust measuring rack is prominently characterized in that a propellant supply pipeline of an engine and a measuring line pipeline are arranged into three segments, namely a vacuum chamber and fixed frame connecting segment, a corrugated pipe connecting segment between a fixed frame and a mobile frame and a mobile frame pipeline and engine pipeline connecting segment. A single working-segment plate spring is adopted to connect the fixed frame with the mobile frame, and three safety limiting devices are arranged. An in-situ thrust calibration device and the safety limiting devices are arranged. Before a test engine is installed and hot rest run is carried out, a working sensor is calibrated. The high-precision liquid or gas rocket engine thrust rack has the advantages that the interference of a pipeline system to engine thrust measurement is greatly reduced; the overall accuracy of the thrust measuring rack is increased; and the damage to the thrust measuring rack due to the disordered engine work can be prevented.

A multi-mode multi-propellant rocket engine capable of operating in a plurality of selected modes.Propellant components may include liquid hydrogen, liquid hydrocarbon, liquid oxygen, liquid fluorine, and liquid air. The liquid oxygen and the liquid air are stored in separate tanks are mixed in a dedicated mixer prior to their injection into the combustion chamber.

An electronic propulsion engine that creates a propulsive force or thrust using electromagnetic forces or electrostatic forces, with an effect that is similar to the thrust of a jet or rocket engine. Forces are generated using electromagnets or capacitor plates that are separated by dielectric spacer cores and are operated with two modulated currents. The two modulated currents are synchronized, but with a relative phase such that the forces on the two magnets or capacitor plates are not balanced. Included are techniques to reduce circuit impedance and control electric-magnetic field dispersion, such as tuned LCR circuits, dielectric core materials between the magnets or capacitor plates, and RF superconductors result in high propulsion efficiencies. The system operates at RF frequencies and can also be used as a communication device.

A remotely-controlled vehicle is disclosed. According to various embodiments, the vehicle includes components that may be added to an existing vehicle or added at the time of manufacture. The components include at least one of a pneumatic projectile launcher and a water cannon, a rocket launcher having one or more rockets launchable from the vehicle, at least one video camera system for capturing and transmitting video images, and a controller for controlling the projectile launcher, water cannon, rocket launcher, and video camera system. Each of the one or more rockets may include at least one solid-propellant rocket motor, and each of the one or more video camera systems may include at least one video camera mounted for selective orientation in at least one plane.

A wet scrubber having a series of chambers for capturing and cooling exhaust gases generated during static test firing of rocket motors. Exhaust gas enters an inlet to a first chamber and is cooled and slowed by a spray solution. HCL gas is condensed and absorbed by the spray solution and precipitates to a liquid slurry at the bottom of the device. The remaining exhaust products enter a demister chamber where nozzles continue to spray the gasses as they pass upward and through a mesh-style demister at the top of the vessel. The demister filters liquid and solid waste particles from the gas stream, and the clean, dry gases are accelerated through a centrifugal fan into the atmosphere. A deflector is positioned within the inlet to the first chamber for containing parts in the event of a motor mal-function.

A weapon system and corresponding ammunition for causing less-than-lethal effects to a target employ a projectile fired from a rifle-mounted, rifle-operated or dedicated rifle-like device and configured for generating less-than-lethal effects to a human target. The projectile includes a rocket motor configured to mitigate velocity decay of the projectile. The rocket motor is preferably a sustainer motor for maintaining a roughly constant cruise velocity of the projectile.

A fuel or fuel blendstocks for jet, gas turbine, rocket, and diesel engines, particularly jet and rocket engines utilizing components of conventional petroleum not currently utilized for jet, gas turbine, rocket, and diesel fuels, such as benzene, butanes, butanes, and methyl tert butyl ether (MTBE) alklyated with aromatic moieties to make monoaromatics used in jet and diesel fuels. Additionally, a fuel having such monoaromatics has multiple desired properties such as higher flash point, low pour point, increased density, better lubricity, aerobic degradability, reduction in toxicity, and additionally can deliver benefits in blendstocks.

A single-stage hypersonic vehicle is comprised of a low-speed and a high-speed propulsion system. The low-speed propulsion system propels the single-stage vehicle to a threshold velocity, after which the high-speed propulsion system then takes over. The low-speed propulsion system includes a combined-cycle engine featuring a swirl generator that is integrated into a turbojet engine to provide a compact turbojet and swirl afterburner-ramjet propulsion system. The high-speed propulsion system includes a hypersonic engine that is operable at the threshold takeover velocity and beyond. In various embodiments, the high-speed propulsion system comprises a scramjet, rocket, or scramjet / rocket engine depending requirements. Benefits of the swirl generator design include its ability to rapidly and efficiently atomize, vaporize, mix and burn the fuel and oxidizer for the low speed propulsion system, significantly reduce length, weight, cooling requirements and complexity for both propulsion systems, while maintaining high propulsion performance and reducing propulsion and launch costs.

Certain embodiments of the invention may include systems, methods, and apparatus for providing a multi-fuel hybrid rocket motor. According to an example embodiment of the invention, a method is provided for producing a multi-fuel hybrid motor. The method can include forming a body, where the body includes one or more intake ports; one or more exit nozzles; one or more channels connecting the one or more intake ports with the one or more exit nozzles; and a plurality of cavities comprising segment walls in communication with the one or more channels. The method also includes depositing a propellant fuel within the plurality of cavities, wherein at least a portion of the propellant fuel is exposed to the one or more channels and wherein the propellant fuel has a higher burn consumption rate than the segment walls.

A hybrid rocket motor is manufactured by photopolymerizing the solid fuel grain in a stereolithography method, wherein fuel grains in a plastic matrix are deposited in layers for building a solid fuel rocket body in three dimensions for improved performance and for a compact design,

A dual expander cycle (DEC) rocket engine with an intermediate closed-cycle heat exchanger is provided. A conventional DEC rocket engine has a closed-cycle heat exchanger terhamlly coupled thereto. The heat exchanger utilizes heat extracted from the engine's fuel circuit to drive the engine's oxidizer turbomachinery.

An independent and conserved source of fuel and / or power comprises a top stage rocket engine firing up to 5000 F. at very high pressures, delivering jet flows up to transonic velocities into a near adiabatic tunnel for mixing in general and / or for transforming reactants introduced to suit specific objectives. The related compression is supplied by an independent prime mover which compresses its exhaust and other recoverable fluids. Low grade flows, thereby upgraded in temperature and pressure, are adiabatically contained, are further upgraded in the tunnel to become part of the prescribed fuel for export at the tunnel ends; or fuel to be fired in a prime mover for electric or other power, or hydrogen for chemical use. Expansion turbines for this purpose are relieved of the load used to compress the excess air in standard gas turbines thus increasing export power. A portion of the expansion turbine's exhaust becomes part of recoverable fluids. When oxygen is used instead of air, the gases through turbines are nitrogen-free with more heat capacity reducing turbine inlet temperatures for the same power. When reactant transformation is specified, the larger water vapor content in the cycle enhances the water gas / shift autothermally for ammonia and / or power and alternatively for pyrolysis cracking for olefins and diolefins. Further, staging rocket engine reactors increases efficiency in boilers and steam turbines; and staging can produce sponge iron and / or iron carbide as well as expansion turbine power and fuel cells for peak and off-peak loads.

The present invention is a space launch system and method to propel a payload bearing craft into earth orbit. The invention has two, or preferably, three stages. The upper stage has rocket engines capable of carrying a payload to orbit and provides the capability of releasably attaching to the lower, or preferably, middle stage. Similar to the lower stage, the middle stage is a reusable booster stage that employs all air breathing engines, is recoverable, and can be turned-around in a short time between missions.

In some implementations 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. In some implementations, 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.

Methods and apparatus are provided for transporting a solid rocket motor, rocket or other object. A transport for an object suitably includes a trailer having a longitudinal axis. A tail support coupled to the trailer has a notch configured to receive a pin affixed to the object to prevent movement along the longitudinal axis during transport. A chock assembly includes a chock and a trolley, wherein the chock is configured to accept the solid rocket motor and to pivot about a rotation axis that is substantially perpendicular to the longitudinal axis of the trailer. By pivoting with the object, the rotable chock maintains intimate contact with the object during transport and / or elevation, thereby improving response to flexing of the transport and reducing the possibility of damage to the object.

The invention relates to a measuring device and a method for detecting thrust vector of a tiny rocket motor used in vacuum. The device for detecting the thrust vector mainly comprises a pressure probe, a movable platform, a rotating table and a data acquisition system. The method comprises the following steps: fixing the pressure probe to the movable platform first, and then adjusting the position of the pressure probe simultaneously, so as to make an axis of the pressure probe parallel to an axis of the tiny rocket motor. When the movable platform moves at constant speed along the direction perpendicular to the axis of the tiny rocket motor, a sensor positioned in the pressure probe senses local dynamic pressure signals in jet current generated by the tiny rocket motor in real time, and the acquisition system acquires and transmits the signals to a data processing and displaying system. Through analyzing symmetry of the dynamic pressure relative to the axis of the tiny rocket motor, the method studies deflection conditions of actual thrust vector and design objective.

A solid propellant rocket engine includes: a propellant formed into a predetermined shape having a combustion cavity and capable of deflagrating into hot, high-pressure gas upon ignition; and a nozzle configured and positioned to expand the hot, high-pressure gas into the atmosphere, where the propellant forms the nozzle as a part of the predetermined shape. A multi-stage rocket unit, a multi-engine rocket stage, and a super-staged rocket each has many rocket engines.

A thermal barrier for extremely high temperature applications consists of a carbon fiber core and one or more layers of braided carbon fibers surrounding the core. The thermal barrier is preferably a large diameter ring, having a relatively small cross-section. The thermal barrier is particularly suited for use as part of a joint structure in solid rocket motor casings to protect low temperature elements such as the primary and secondary elastomeric O-ring seals therein from high temperature gases of the rocket motor. The thermal barrier exhibits adequate porosity to allow pressure to reach the radially outward disposed O-ring seals allowing them to seat and perform the primary sealing function. The thermal barrier is disposed in a cavity or groove in the casing joint, between the hot propulsion gases interior of the rocket motor and primary and secondary O-ring seals. The characteristics of the thermal barrier may be enhanced in different applications by the inclusion of certain compounds in the casing joint, by the inclusion of RTV sealant or similar materials at the site of the thermal barrier, and / or by the incorporation of a metal core or plurality of metal braids within the carbon braid in the thermal barrier structure.

The invention relates to a method for vertically processing a cooling passage of a liquid rocket engine thrust jet, which belongs to the processing method field of liquid rocket engine thrust jet. In the invention, the thrust jet adopts a vertical clamp; a single-point laser sensor scans bus bars strip by strip to acquire the actual profile of the thrust jet and a sampled data set to reconstruct the bus bars of the thrust jet strip by strip; the buses of the sampled data set are smoothed strip by strip to acquire a thrust jet smoothed data set; the data set of all the bus bars are compressed to acquire a compressed data set; and a cubic spline is used for performing interpolation to approximate the actual bus bars of the thrust jet. The milling tool position is calculated by dual milling head for symmetrical milling and high speed back chipping. The method realizes the integration of measurement and processing, finishes the treatments of profile measurement, milling groove and back chipping by one clamping, can satisfy the processing requirements on different physical dimensions and different bus bar linetypes, has high processing efficiency and good consistency, and realizes the digital processing with high efficiency and high precision of the cooling passage of the liquid rocket engine thrust jet.

Embodiments include active protection systems and methods for an aerial platform. An onboard system includes one or more radar modules, detects aerial vehicles within a threat range of the aerial platform, and determines if any of the plurality of aerial vehicles are an aerial threat. The onboard system also determines an intercept vector to the aerial threat, communicates the intercept vector to an eject vehicle, and causes the eject vehicle to be ejected from the aerial platform to intercept the aerial threat. The eject vehicle includes a rocket motor to accelerate the eject vehicle along an intercept vector, alignment thrusters to rotate a longitudinal axis of the eject vehicle to substantially align with the intercept vector, and divert thrusters to divert the eject vehicle in a direction substantially perpendicular to the intercept vector. The eject vehicle activates at least one of the alignment thrusters responsive to the intercept vector.

The present invention provides a transpiration cooled rocket motor which operates at low combustion pressures, inherently provides stabilized combustion, and operates over a desired range of L-star. In a preferred embodiment the inventive transpiration cooled rocket motor includes: a housing; a porous injector sleeve; a propellant injector; and a nozzle. Preferably the injector sleeve lines the inside walls of the housing such that the inside volume of the sleeve forms the chamber for the motor. Liquid fuel passes inwardly through the pores of the injector sleeve and is injected from the sleeve into the chamber, thus cooling the walls of the rocket motor.

A high-efficiency spacecraft propulsion system, including includes electric pumps inserted in the oxidizer and fuel lines that increase liquid apogee engine (“LAE”) operating pressure and reduce tank-operating pressure. An on-board computer generates pump drive signals, in response to measured oxidizer and fuel line pressures, that are input to the pump controller electronics. The controller electronics provides current drives to the pump motors. The system uses an LAE that can operate at higher thrust chamber pressures (e.g., 500 psia) than standard LAEs, where pump-fed rocket motors have integrated turbopumps that are fuel operated. The turbopump increases LAE complexity and cost, reduces fuel efficiency, and is not compatible with active thrust and mixture ratio control.

The invention discloses a mobile test platform for low-thrust rocket engines. The mobile test platform comprises a mobile platform, a running support, a propellant feeding system and a gas distribution system. The running support, the propellant feeding system and the gas distribution system are integrally mounted on the mobile platform. The mobile platform comprises a testing rack mounting structure, wheels, a load-carrying structure, an inner cavity and a gas distribution panel. The testing rack comprises a motionless support, plate springs, a movable support, an engine support, a thrust support and a thrust sensor. The propellant feeding system comprises a fuel supply system and an oxidant supply system. The gas distribution system comprises a high-pressure gas tank, a pressurizing gas pipeline and a blowdown gas pipeline. All system components are integrated on one mobile platform, so that the mobile test platform has the advantages of miniaturization, portability and the like.

An ignition and chemical reaction accelerator and flame stabilizer and a high-performance speed-type internal combustion engine using this device are provided, whereby ignition and the spreading and holding of flames can be dramatically improved in a gas turbine, a ram machine, a rocket engine, or another speed-type internal combustion engine. An ignition and chemical reaction accelerator and flame stabilizer of a speed-type internal combustion engine comprises a spark plug (110) for preparing charged particles in a predetermined location in a combustor (104) of the speed-type internal combustion engine, and a microwave oscillator (116) and antenna (112) for inducing plasma with a working fluid in the combustor as a starter material by irradiating the charged particles and their surrounding vicinity with microwave pulses; wherein a region in which sufficient conditions for performing combustion are met is formed in the combustor by supplying an active chemical species produced from the working fluid by the effect of the plasma.

A projectile comprises an imaging seeker at a front of the projectile; a front warhead behind the imaging seeker; a power supply; an electronics unit connected to the power supply and comprising a microprocessor circuit board, a voltage regulator circuit board, an inertial measurement circuit board and a fuze and safe and arm circuit board, all electrically connected to each other, the microprocessor circuit board also being connected to the imaging seeker; a rear warhead, the front and rear warheads being electrically connected to the safe and arm circuit board; a rocket motor electrically connected to the electronics unit; foldable fins mounted at the rear of the projectile; a shell that encases the front warhead, the power supply, the electronics unit, the rear warhead and the rocket motor; and a maneuver mechanism disposed in the shell and electrically connected to the microprocessor circuit board.