A solid rocket motor uses at least one thermally conductive wire or at least one pair of electrically conductive wires to increase a burn surface area of a propellant grain and thus a thrust of the rocket motor. The rocket motor includes a pulse chamber containing a burnable propellant grain, a propellant inhibited center bore bonded to surfaces of the burnable propellant grain, and at least one conductive wire coupled to the burnable propellant grain and arranged in variable regions along the propellant inhibited center bore. The conductive wire is configured for passive or active activation to ignite the propellant inhibited center bore that subsequently burns in the variable regions. The thermally conductive wire is formed of a refractory metal or refractory alloy material that enables the entire length of the wire to be heated simultaneously or nearly simultaneously when the wire is passively activated.

A solid rocket motor uses at least one thermally conductive wire or at least one pair of electrically conductive wires to increase a burn surface area of a propellant grain and thus a thrust of the rocket motor. The rocket motor includes a pulse chamber containing a burnable propellant grain, a propellant inhibited center bore bonded to surfaces of the burnable propellant grain, and at least one conductive wire coupled to the burnable propellant grain and arranged in variable regions along the propellant inhibited center bore. The conductive wire is configured for passive or active activation to ignite the propellant inhibited center bore that subsequently burns in the variable regions. The thermally conductive wire is formed of a refractory metal or refractory alloy material that enables the entire length of the wire to be heated simultaneously or nearly simultaneously when the wire is passively activated.

A solid rocket motor uses at least one thermally conductive wire or at least one pair of electrically conductive wires to increase a burn surface area of a propellant grain and thus a thrust of the rocket motor. The rocket motor includes a pulse chamber containing a burnable propellant grain, a propellant inhibited center bore bonded to surfaces of the burnable propellant grain, and at least one conductive wire coupled to the burnable propellant grain and arranged in variable regions along the propellant inhibited center bore. The conductive wire is configured for passive or active activation to ignite the propellant inhibited center bore that subsequently burns in the variable regions. The thermally conductive wire is formed of a refractory metal or refractory alloy material that enables the entire length of the wire to be heated simultaneously or nearly simultaneously when the wire is passively activated.

A solid rocket motor uses at least one thermally conductive wire or at least one pair of electrically conductive wires to increase a burn surface area of a propellant grain and thus a thrust of the rocket motor. The rocket motor includes a pulse chamber containing a burnable propellant grain, a propellant inhibited center bore bonded to surfaces of the burnable propellant grain, and at least one conductive wire coupled to the burnable propellant grain and arranged in variable regions along the propellant inhibited center bore. The conductive wire is configured for passive or active activation to ignite the propellant inhibited center bore that subsequently burns in the variable regions. The thermally conductive wire is formed of a refractory metal or refractory alloy material that enables the entire length of the wire to be heated simultaneously or nearly simultaneously when the wire is passively activated.

Propulsion systems that generate thrust from pressure generated by thermally decomposing a chemical blowing agent (CBA). In some embodiments, the CBA decomposes exothermically such that once thermal decomposition has been initiated, the thermal decomposition continues without additional energy input. In some embodiments, the CBA is utilized in a digital-microthruster array containing microthrusters that can be individually activated to provide thrust. In some embodiments, a CBA may be stored in one or more CBA-storage chambers that can be individually activated to charge and/or recharge a pressure tank that stores gas from the CBA decomposition under pressure for providing thrust. These and other embodiments are disclosed. Such propulsion systems can be used for any of a variety of spacecraft, including micro- and nano-satellites. Corresponding methods of generating thrust are also disclosed.

A rocket engine has: a combustion chamber having a chamber inlet for receiving an oxidizer and a chamber outlet for expelling combustion gases in an environment outside the combustion chamber; a manifold having a manifold inlet fluidly connectable to a source of the oxidizer and a manifold outlet; a catalyst having a catalyst inlet fluidly connected to the manifold outlet and a catalyst outlet; and an injector plate having a injector inlet fluidly connected to the catalyst outlet and an injector outlet fluidly connected to the chamber inlet.

A hybrid rocket motor includes a solid fuel element, and an oxidizer tank containing an oxidizer. The solid fuel element adjoins and at least partially defines a combustion chamber in which the solid fuel and the oxidizer are burned, to produce thrust from the hybrid rocket motor. The oxidizer tank is at least partially within the combustion chamber, and the entire oxidizer tank may be within the combustion chamber. The oxidizer tank may be protected by an insulating material, which may also serve as a structural material that contains the pressure of the oxidizer. The insulating material and the fuel material may both be polymer-based materials, although they may be different materials having different characteristics, for example including different additives to the same polymer material. The fuel element and the oxidizer tank may be made by additive manufacturing processes, for example by adding different materials in different locations.

A hybrid rocket motor includes a solid fuel element, and an oxidizer tank containing an oxidizer. The solid fuel element adjoins and at least partially defines a combustion chamber in which the solid fuel and the oxidizer are burned, to produce thrust from the hybrid rocket motor. The oxidizer tank is at least partially within the combustion chamber, and the entire oxidizer tank may be within the combustion chamber. The oxidizer tank may be protected by an insulating material, which may also serve as a structural material that contains the pressure of the oxidizer. The insulating material and the fuel material may both be polymer-based materials, although they may be different materials having different characteristics, for example including different additives to the same polymer material. The fuel element and the oxidizer tank may be made by additive manufacturing processes, for example by adding different materials in different locations.

An insulation material includes a matrix comprising a reaction product formed from a silicon carbide precursor resin and a silicon dioxide precursor resin. At least one filler, such as hollow glass microspheres and/or carbon fiber is dispersed within the matrix. A rocket motor includes a case, the insulation material within and bonded to the case, and a solid propellant within the case. An insulation precursor includes a silicon carbide precursor resin, a silicon dioxide precursor resin, and the at least one filler. Related methods are also disclosed.

An insulation material includes a matrix comprising a reaction product formed from a silicon carbide precursor resin and a silicon dioxide precursor resin. At least one filler, such as hollow glass microspheres and/or carbon fiber is dispersed within the matrix. A rocket motor includes a case, the insulation material within and bonded to the case, and a solid propellant within the case. An insulation precursor includes a silicon carbide precursor resin, a silicon dioxide precursor resin, and the at least one filler. Related methods are also disclosed.