Abstract: A rocket engine has a combustion chamber having an inlet and an outlet, the inlet fluidly connectable to a source of oxidizer, the outlet in fluid communication with an environment outside the combustion chamber for expelling combustion gases, a first fuel having a first solid propellant and a second fuel having a second solid propellant, the first and second fuels located within the combustion chamber and configured to be exposed to the oxidizer injected in the combustion chamber via the inlet, the first solid propellant having a regression rate greater than that of the second solid propellant.

Abstract: Various embodiments of a vortex hybrid motor are described herein. In some embodiments, the vortex hybrid motor may include a combustion zone defined by a fuel core and/or motor housing. The combustion zone may include an upper zone and a central zone that each contribute to thrust created by the vortex hybrid motor. In some embodiments, an injection port configuration is described that includes a proximal injection port that may be controlled for modulating a delivery of an amount of oxidizer for adjusting an oxidizer-to-fuel ratio. In some embodiments, a fuel core configuration is described that provides radially varying gradients of fuel in order to achieve desired thrust profiles. In some embodiments, the fuel core may include a support structure and/or a proximal end of a nozzle of the vortex hybrid motor may extend into the fuel core.

Abstract: A hybrid propulsion system for a small satellite package consisting of a main rocket motor containing a solid propellant with multiple oxidizer tanks positioned to direct oxidizer into the rocker motor, thereby producing a desired thrust necessary for orbit insertion and/or orbit correction. Additionally, oxidizers can serve a dual function in controlling cold fuel thrusters for attitude adjustment.

Abstract: A hybrid rocket includes a housing having first and second ends, a solid-grain fuel material in the housing and defining a bore extending from end to end, two electrodes positioned adjacent to the fuel material to ignite the fuel material at the first end, a primary oxidizer port positioned at the first end to inject a primary oxidizer to flow in a downstream direction from the first end to the second end, a nozzle positioned at the second end and having a converging portion and a diverging portion, and a secondary oxidizer port to inject a secondary oxidizer downstream of the converging portion. The bore has a geometry configured to produce a hot-gas, fuel-rich mixture at the nozzle as the fuel material and primary oxidizer burn while flowing downstream. The diverging portion of the nozzle is configured to spontaneously combust the secondary oxidizer and the hot-gas, fuel-rich mixture.

Abstract: A motor has an oxidizer injector, the oxidizer injector is mainly suitable for using in a combustion chamber, the oxidizer injector has a body having a first runner assembly and a second runner assembly arranged along an axis, the first runner assembly injects oxidizer into the combustion chamber to form a forward swirl, and the second runner assembly injects oxidizer into the combustion chamber to form a reverse swirl, the axial torsion generated by the forward swirl and the axial torsion generated by the reverse swirl counteract each other, so as to solve the problem of axial torsion imbalance in the combustion chamber.

Abstract: A rocket engine with an improved solid fuel segment mainly comprises a combustion chamber, a solid fuel segment installed in the combustion chamber, and an oxidizer injector installed at one end of the combustion chamber. The solid fuel segment surrounds and forms a trajectory to allow the oxidizer injector to inject oxidizer into the trajectory, in particular, on the solid fuel segment is formed with a plurality of protrusions, between the each two protrusions are defined a recess, a flame holding hot-gas region is formed between the protrusion and the recess, so as to produce eddy current in the flame holding hot-gas region when the propellant mixture is burned inside the trajectory, such that the whole solid fuel segment can produce even regression rate and high combustion efficiency.

Abstract: An injection device, a combustor, and a rocket engine include a device main body partitioned into a fuel manifold and an oxidant manifold, and a plurality of injectors arranged at predetermined intervals in the device main body to inject fuel and oxidant into a combustion chamber, each of the injectors includes a LOx channel including a proximal end portion communicating with the oxidant manifold and a distal end portion communicating with the combustion chamber, a restrictor provided on the proximal end portion of the LOx channel and a GH2 channel including a proximal end portion communicating with the fuel manifold and a distal end portion communicating with the combustion chamber, and the restrictors have different shapes.

Abstract: A hybrid rocket solid fuel grain having a cylindrical shape and defining a center port is additive manufactured from a compound of thermoplastic fuel and passivated nanocomposite aluminum additive. The fuel grain comprises a stack of fused layers, each layer formed as a plurality of fused abutting concentric circular beaded structures arrayed to define a center port. During operation, an oxidizer is introduced along the center port, with combustion occurring along the exposed port wall. Each circular beaded structure defines geometry that increases the surface area available for combustion. As each layer ablates the next abutting layer, exhibiting a similar geometry, is revealed, undergoes a gas phase change, and ablates. This process repeats and persists until oxidizer flow is terminated or the fuel grain material is exhausted. To safely achieve this construction, a fused deposition additive manufacturing apparatus, modified to shield the nanocomposite material from the atmosphere, is used.

Abstract: An anaerobic propulsion device of hybrid type. The propulsion device includes a combustion chamber into which an oxidant is injected thereto comes into contact with a fuel. The chamber includes an orifice, notably a nozzle, to inject the combustion gases. The fuel takes the form of solids divided into granules, typically grains, powder or beads. The propulsion device further includes a segregation device arranged between a reservoir for granules of solid fuel, and the combustion chamber. The segregation device is configured to prevent granules of fuel in solid form from passing between the reservoir and the combustion chamber. The segregation device includes orifices to allow the passage of the fuel, once the fuel has become pasty, liquid or vaporized.

Abstract: An additively manufactured solid fuel grain for a hybrid rocket engine having a cylindrical shape, defining a center combustion port and comprising a stack of fused layers of polymeric material suitable for hybrid rocket fuel. Each layer is formed as a plurality of fused abutting concentric beads of solidified material arrayed around the center port. An oxidizer is introduced into the solid fuel grain through the center port, with combustion occurring along the exposed surface area of the solid fuel grain center port wall. Each concentric bead possesses a surface pattern that increases the combustion surface area and when stacked forms a rifling pattern of undulations that induces oxidizer-fuel gas axial flow to improve combustion efficiency. The port wall surface pattern persists during the rocket engine's operation as the fuel phase changes from solid to gas and is ablated.

Abstract: A hybrid rocket solid fuel grain having a cylindrical shape and defining a center port is additive manufactured from a compound of thermoplastic fuel and passivated nanocomposite aluminum additive. The fuel grain comprises a stack of fused layers, each formed as a plurality of fused abutting concentric circular beaded structures of different radii arrayed defining a center port. During operation, an oxidizer is introduced along the center port, with combustion occurring along the exposed port wall. Each circular beaded structure possesses geometry that increases the surface area available for combustion. As each layer ablates the next abutting layer, exhibiting a similar geometry is revealed, undergoes a gas phase change, and ablates. This process repeats and persists until oxidizer flow is terminated or the fuel grain material is exhausted. To safety achieve this construction, a fused deposition additive manufacturing apparatus, modified to shield the nanocomposite material from the atmosphere is used.

Abstract: Ceramic catalyst carriers that are mechanically, thermally and chemically stable in a ionic salt monopropellant decomposition environment and high temperature catalysts for decomposition of liquid high-energy-density monopropellants are disclosed. The ceramic catalyst carrier has excellent thermal shock resistance, good compatibility with the active metal coating and metal coating deposition processes, melting point above 1800° C., chemical resistance to steam, nitrogen oxides and acids, resistance to sintering to prevent void formation, and the absence of phase transition associated with volumetric changes at temperatures up to and beyond 1800° C.

Abstract: Propulsion systems and method for making a propulsion system include additively manufacturing a casing body into a single-piece structure having no bonded or bolted joints. The casing body defines a combustion chamber therein and is at least partially composed of a material useful as a solid rocket fuel and capable of being consumed during combustion. Other embodiments are also described.

Abstract: Supply of a liquid component in a combustion chamber of a rocket engine is controlled by a feed valve provided with slide valve mobile between an open position and a closed position of at least one supply pipe, which has an inlet that communicates with a tank for containing the liquid component and an outlet that communicates with the combustion chamber. The displacement of the slide valve from its closed position to its open position is triggered by a pressurized fluid supplied to the outlet of the supply pipe.

Abstract: 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,

Abstract: An igniter for a rocket engine or motor comprising a combustion chamber with a solid fuel, an inlet for supplying an oxidizer to the combustion chamber to ignite the solid fuel and an outlet for discharging exhaust gas, wherein the igniter is arranged to discharge exhaust gas to the rocket engine for igniting the rocket engine. The igniter can be used for multiple ignitions and can also be re-used after re-filling.

Abstract: 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, the hybrid rocket motor including radial channels for defining a desired burn profile including the oxidizer to fuel burn ratio.

Abstract: The present invention relates to improved rocket engine systems. In one embodiment, an improved rocket engine system includes a propellant source, at least one power source, at least one power source motor, a rocket engine, and at least one pump. The improved rocket engine system may further include at least one of the following: at least one controller, at least one propellant valve, and a propellant pressurizing source.

Abstract: 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.

Abstract: The Multiple Use Plug Hybrid (for) Nanosats (“MUPHyN”) prototype thruster is being developed to fill a niche application for NanoSat scale spacecraft propulsion. The MUPHyN thruster uses safe-handling and inexpensive nitrous oxide (N20) and acrylonitrile-butadiene-styrene (ABS) as propellants. The MUPHyN thruster can provide an enhanced propulsive capability that will enable multiple NanoSats to be independently repositioned after deployment from the parent launch vehicle. Because the environmentally benign propellants are mixed only within the combustion chamber once the ignition is initiated, the system is inherently safe and can be piggy-backed on a secondary payload with no overall mission risk increase to the primary payload.

Abstract: 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, the hybrid rocket motor including buried radial channels for defining a desired burn profile including the oxidizer to fuel burn ratio.

Abstract: A liquid propellant tank for storing a liquid propellant A and supplying vapor produced by evaporation of part of the liquid propellant A to an external location comprises a tank body for storing the liquid propellant A, and a plurality of holder plates arranged inside the tank body, around an axis L of the tank body. The holder plates cause the liquid propellant A to adhere to them by surface tension, thereby establishing, inside the tank body, a liquid propellant holding area LA in which the liquid propellant A is held and a gas accumulation area GA in which vapor produced by evaporation of part of the liquid propellant A accumulates. The tank body has a propellant inlet open into the liquid propellant holding area LA and a gas outlet open into the gas accumulation area GA.

Abstract: A thruster designed to produce thrust by introducing a propellant to a catalyst layer from a propellant valve via a propellant introduction member and spurting out a gas resulting from decomposition of the propellant occurring on the catalyst layer, wherein the thruster comprises a chamber of an Ni alloy for holding the catalyst layer inside, the propellant introduction member is made of an Ni alloy and connects the propellant valve and the chamber, and a propellant valve flange of a Ti alloy with a plurality of columns of a Ti alloy is arranged between the chamber and the propellant valve to support the chamber by the columns.

Abstract: An apparatus for igniting a larger rocket motor is provided. The apparatus may be a smaller rocket motor that can be ignited hypergolically, when a pressurized oxidizer contacts hypergolic fuel grains of the smaller rocket motor. The hypergolic ignition causes the larger rocket motor to be ignited. The hypergolic ignition of the smaller rocket motor may be stopped after the larger rocket motor is ignited, and the remaining hypergolic fuel grains and the pressurized oxidizer can be reserved for reigniting the larger rocket motor at a later time.

Abstract: A rocket engine system with a fuel conversion system in communication with a gas generator.

Abstract: The present invention describes a hybrid rocket motor that includes a first solid reactant and at least one thrust nozzle and at least one moveable combustion control member within the hybrid rocket motor that restricts the contact of a first fluid reactant in the combustion chamber. In this way, it is then possible to regulate the exposure of the solid reactant to the fluid reactant and thus control thrust.

Abstract: One subject of the present invention is a propulsion method comprising: the injection, into at least one combustion chamber (1), of at least one liquid oxidizer (OX) and of hydrogen (H2); the combustion of said at least one liquid oxidizer (OX) and hydrogen (H2), in said at least one combustion chamber (1), for the generation of combustion gases; and expulsion of said combustion gases. Said process comprises, upstream of said injection: the generation of at least one portion of said hydrogen (H2), advantageously of all said hydrogen (H2), from at least one solid compound (5?); said generation from said at least one solid compound (5?) comprising a combustion reaction between said at least one solid compound (5?) chosen from alkali metal borohydrides, alkaline-earth metal borohydrides, borazane, polyaminoboranes and mixtures thereof and an oxidizing charge (5?). Another subject of the present invention is a propulsion device suitable for the implementation of said method.

Abstract: An apparatus and method to electrocatalytically induce the decomposition of a liquid propellant for propulsion. This is accomplished by providing a reaction chamber filled with catalyst that has a first electrode, such as near the center, and the other electrode disposed away from the first electrode. A charge source that is capable of applying voltage is connected to the electrodes. When the reaction chamber is dosed with a propellant such as an aqueous based amine propellant, such as HAN, an electric current flows between the electrodes and the propellant. This initiates the decomposition of the propellant which is driven to completion by the catalyst bed in order to be used for power generation.

Abstract: A solid rocket motor includes a combustion chamber bounded by an outer casing, a propellant grain within the combustion chamber, and an igniter within the outer casing for igniting the propellant grain. A nozzle is coupled to the combustion chamber for releasing hot gasses evolved from burning the propellant grain to provide thrust for propelling the solid rocket motor. The propellant grain is a self-extinguishing propellant grain that includes at least one fuel, at least one oxidizing agent, at least one binder, and at least one surfactant that imparts the self-extinguishing property. The propellant grain provides a burning rate as a function of pressure that includes a negative pressure dependence portion, wherein the burning rate in the negative pressure dependence portion decreases with increasing pressure until a cutoff pressure is reached which results in extinguishment of the propellant grain.

Abstract: Control and/or drive device for a flying body for ejecting hot gas streams of a combusted fuel combination of at least a first and second component. Device includes a first hollow chamber body structured and arranged to contain first component, a second hollow chamber body structured and arranged to contain second component, a controllable fuel valve arranged between first hollow chamber body and second hollow chamber body to control feed of first component to second hollow chamber body, and a plurality of outlets structured and arranged to eject respective hot gas streams for influencing a flight path of flying body. Second hollow chamber body is formed as a combustion chamber for combusting the at least first and second components within second hollow chamber body to generate respective hot gas streams, and plurality of outlets are connected to the second hollow chamber body.

Abstract: A motor includes an annular solid fuel between an inside diameter and an outside diameter. The solid fuel has a series of radial notches that define segments of fuel between them. The notches allow for faster burning of the fuel, while still allowing structural integrity of the fuel segments to be retained during the burning process.

Abstract: Disclosed is a propulsion system having a structural configuration that provides easy and convenient access to the interior regions of a liquid fuel tank and a hybrid rocket motor case. In one embodiment, the propulsion system comprises: a hybrid rocket motor case and a fuel tank coupled to the hybrid motor case. The motor case is configured to hold solid rocket fuel and the fuel tank defines an internal volume configured to hold a fluid oxidizer. A bulkhead is interposed between the motor case and the fuel tank, wherein at least one access passageway extends through the bulkhead. The access passageway provides exterior access to the interior volume of the motor case or the internal volume of the storage tank while the hybrid rocket motor is coupled to the fuel tank.

Abstract: Disclosed is a propulsion system having a structural configuration that provides easy and convenient access to the interior regions of a liquid fuel tank and a hybrid rocket motor case. In one embodiment, the propulsion system comprises: a hybrid rocket motor case and a fuel tank coupled to the hybrid motor case. The motor case is configured to hold solid rocket fuel and the fuel tank defines an internal volume configured to hold a fluid oxidizer. A bulkhead is interposed between the motor case and the fuel tank, wherein at least one access passageway extends through the bulkhead. The access passageway provides exterior access to the interior volume of the motor case or the internal volume of the storage tank while the hybrid rocket motor is coupled to the fuel tank.

Abstract: An insensitive combined cycle missile propulsion system includes a solid fuel contained within a first section of the missile, a liquid oxidizer contained within a second section of the missile and a solid oxidizer contained within a third section of said missile. A first conduit has a first valve communicating the fuel and the oxidizer and a second conduit, spatially removed from the first conduit, has a second valve communicating the fuel and the oxidizer. An inlet system for delivering atmospheric oxygen for combustion with the fuel rich gases generated within the missile and a nozzle exhausts combustion products that result from combustion of the fuel, the liquid and solid oxidizers, and air.

Abstract: Propellant management systems and methods are provided for controlling the delivery of liquid propellants in a space launch vehicle utilizing multiple rockets. The propellant management systems and methods may be configured to enable substantial simultaneous depletion of liquid propellants in each of a plurality of active rockets during operation of various booster stages of the launch vehicle.

Abstract: The present invention provides a motor including a combustion chamber, an oxidizer injector, a vortex generators and a nozzle. The combustion chamber can be used to dispose a solid fuel, and the oxidizer injector is used to control the flow rate of an oxidizer and to inject the oxidizer into the combustion chamber. The vortex generators is disposed on the inner wall of the combustion chamber for generating eddies to enhance the mixing of the fuel and the oxidizer. Additionally, the nozzle is connected with the combustion chamber for exhausting the gas generated by the combustion of the propellants.

Abstract: A gas delivery system includes a gas booster module for delivering natural gas from a utility gas service to power generation equipment installed in or around a building in a manner that meets the minimum volume and pressure requirements of the power generation equipment. The gas delivery system advantageously uses pipe of a relatively small size for delivering gas to the power generation equipment, thereby substantially reducing installation costs and eliminating the need for a welded gas line. The gas delivery system also provides a control system that facilitates close control over the gas flow and ensures compliance with local building codes and safety regulations and requirements.

Abstract: Provided is a hybrid rocket using catalytic decomposition of oxidizer, which generates thrust by injecting high temperature oxygen and steam formed by catalytic decomposing liquid oxidizer into solid fuel and thereby auto igniting and combusting the solid fuel without a separate igniter.

Abstract: A hybrid rocket motor is formed from a single piece of material, such as, for example, by extrusion or injection molding. The rocket motor includes various components, such as casing, structure, oxidizer tank, combustion chamber, fuel, port and nozzle that are all formed from a single piece of material. The material can be, for example, a material that can be used as solid rocket fuel.

Abstract: A bi-propellant rocket motor having controlled thermal management is disclosed. The rocket motor produces thrust using a solid or gel-phase primary propellant that can be either fuel- or oxidizer-rich, with a complementary self-pressurizing secondary propellant selected to balance the primary propellant in terms of the equivalence ratio. The motor houses multiple propellant grains arranged in such a configuration that each chamber containing a primary grain serves as both propellant storage and the main combustion chamber for that propellant grain as it burns with the secondary propellant. The secondary propellant is stored separately, and the flow routed past the primary propellant chamber to provide cooling for adjacent primary propellant chambers limiting the temperature rise in the motor structure.

Abstract: Activation of a propellant in a constant volume container causes a phase change material to rapidly expand so that the pressure in the container increases. Programmability and sequential actuation are enabled by patterning the phase change material into the integrated device. The pressure generated may be used to activate an energy transducer such as a high pressure turbine, a piezoelectric material, and an elastic strain material. This provides a hybrid actuation system of electrical energy, pneumatic and hydraulic power. The pressure change in a constant volume container is also harnessed to provide a microbattery.

Abstract: The present invention provides a heretofore-unknown use for zirconium nitride as a hydrogen peroxide compatible protective coating that was discovered to be useful to protect components that catalyze the decomposition of hydrogen peroxide or corrode when exposed to hydrogen peroxide. A zirconium nitride coating of the invention may be applied to a variety of substrates (e.g., metals) using art-recognized techniques, such as plasma vapor deposition. The present invention further provides components and articles of manufacture having hydrogen peroxide compatibility, particularly components for use in aerospace and industrial manufacturing applications. The zirconium nitride barrier coating of the invention provides protection from corrosion by reaction with hydrogen peroxide, as well as prevention of hydrogen peroxide decomposition.

Abstract: A fixed sized bell rocket nozzle is lined with a layer of combustible material that is ignited during launch ignition and burns to outgas into the rocket exhaust for spatially variably confining the exhaust and perfecting an effective variably sized altitude compensating exhaust nozzle that maximizes lift during the launch of a spacecraft into orbit.

Abstract: An ultra-compact aerovortical swirl-enhanced combustion (ASC) system features an aerovortical swirl generator for use in rocket thrusters utilizing hypergolic or non-hypergolic propellants. The ACS thruster can be sized for diameters ranging from about 0.5 to about 2.0 inches, and producing thrust levels of approximately 5 lbf to about 250 lbf. A plurality of helicoid flow channels in the swirl generator introduces swirl into a flow stream of a first propellant within ultra-compact sized rocket thrusters. The ASC system also includes injectors for introducing a second liquid propellant into the swirling flowfield to promote rapid and efficient atomization, mixing and vigorous combustion, which, results in major improvements in combustion and propulsion performance over current rocket thrusters, but in much shorter combustor systems.

Abstract: Control and/or drive device for a flying body for ejecting hot gas streams of a combusted fuel combination of at least a first and second component. Device includes a first hollow chamber body structured and arranged to contain first component, a second hollow chamber body structured and arranged to contain second component, a controllable fuel valve arranged between first hollow chamber body and second hollow chamber body to control feed of first component to second hollow chamber body, and a plurality of outlets structured and arranged to eject respective hot gas streams for influencing a flight path of flying body. Second hollow chamber body is formed as a combustion chamber for combusting the at least first and second components within second hollow chamber body to generate respective hot gas streams, and plurality of outlets are connected to the second hollow chamber body.

Abstract: A rocket engine includes a combustor assembly for carrying out a combustion process of fuel and oxidizer rocket propellants to produce thrust. A swirl generator is positioned within the combustor assembly to produce a turbulent flowfield of the fuel and oxidizer rocket propellants within the combustor assembly.

Abstract: 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, the hybrid rocket motor including radial channels for defining a desired burn profile including the oxidizer to fuel burn ratio.

Abstract: 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,

Abstract: A hybrid rocket solid fuel grain having a cylindrical shape and defining a center port is additive manufactured from a compound of thermoplastic fuel and passivated nanocomposite aluminum additive. The fuel grain comprises a stack of fused layers, each layer formed as a plurality of fused abutting concentric circular beaded structures arrayed to define a center port. During operation, an oxidizer is introduced along the center port, with combustion occurring along the exposed port wall. Each circular beaded structure defines geometry that increases the surface area available for combustion. As each layer ablates the next abutting layer, exhibiting a similar geometry, is revealed, undergoes a gas phase change, and ablates. This process repeats and persists until oxidizer flow is terminated or the fuel grain material is exhausted. To safely achieve this construction, a fused deposition additive manufacturing apparatus, modified to shield the nanocomposite material from the atmosphere, is used.

Abstract: A hybrid rocket solid fuel grain having a cylindrical shape and defining a center port is additive manufactured from a compound of thermoplastic fuel and passivated nanocomposite aluminum additive. The fuel grain comprises a stack of fused layers, each layer formed as a plurality of fused abutting concentric circular beaded structures arrayed to define a center port. During operation, an oxidizer is introduced along the center port, with combustion occurring along the exposed port wall. Each circular beaded structure defines geometry that increases the surface area available for combustion. As each layer ablates the next abutting layer, exhibiting a similar geometry, is revealed, undergoes a gas phase change, and ablates. This process repeats and persists until oxidizer flow is terminated or the fuel grain material is exhausted. To safely achieve this construction, a fused deposition additive manufacturing apparatus, modified to shield the nanocomposite material from the atmosphere, is used.