A nuclear thermal propulsion rocket engine. A source of fissionable material is provided in a bed of fuel pebbles located in a reactor. A fluid having neutron moderating constituents, such as hydrogen and/or carbon, therein, is provided, which may be in the form of methane, or ethane, or a combination thereof, or may further include various isotopes of hydrogen. An external neutron source is provided using a neutron beam generator. Reactor design geometry provides containment of fissionable material, and for any byproducts of fission reactions, in the reactor during acceleration of the rocket. Impact of neutrons on fissionable material results in a nuclear fission reaction conditions in the reactor, resulting in release of heat energy to fluids provided to the reactor. The reactor is sized and shaped to contain fuel pebbles containing fissionable material, and to confine expandable fluids as they remove heat from fuel pebbles. the heated fluids are discharged out through a throat, into a rocket engine expansion nozzle for propulsive discharge, The design provides a rocket engine with a specific impulse in the range of from about eight hundred (800) seconds to about twenty five hundred (2500) seconds.

A nuclear thermal propulsion rocket engine. A source of fissionable material is provided in a bed of fuel pebbles located in a reactor. A fluid having neutron moderating constituents, such as hydrogen and/or carbon, therein, is provided, which may be in the form of methane, or ethane, or a combination thereof, or may further include various isotopes of hydrogen. An external neutron source is provided using a neutron beam generator. Reactor design geometry provides containment of fissionable material, and for any byproducts of fission reactions, in the reactor during acceleration of the rocket. Impact of neutrons on fissionable material results in a nuclear fission reaction conditions in the reactor, resulting in release of heat energy to fluids provided to the reactor. The reactor is sized and shaped to contain fuel pebbles containing fissionable material, and to confine expandable fluids as they remove heat from fuel pebbles. the heated fluids are discharged out through a throat, into a rocket engine expansion nozzle for propulsive discharge, The design provides a rocket engine with a specific impulse in the range of from about eight hundred (800) seconds to about twenty five hundred (2500) seconds.

A multipulse rocket motor includes a secondary pulse, fired after a primary pulse of the motor, that includes a thermal insulator having channels therein, around a propellant grain of the secondary pulse. The channels provide a way to equalize pressure on the propellant grain of the secondary pulse, to reduce stresses on the propellant grain as the primary pulse is operating. The channels may extend along most or substantially all of a length of the secondary pulse. The channels may be defined by material strips of thermal insulator material evenly circumferentially spaced around the secondary pulse.

A rocket throat insert including an annular body having a radially inner annular wall portion and a radially outer annular portion. The inner wall portion has a contoured radially inner surface defining a nozzle throat. The outer portion includes an annular buttressing structure supporting the inner wall portion and defining one or more insulation gaps arranged annularly around the inner wall portion. The insulation gaps restrict the radial flow of heat through the annular body.

An insulated blast tube includes an insulating layer of a burn resistant material such as phenolic resin formed on an interior surface of the blast tube to provide the necessary erosion and thermal insulation properties to protect the blast tube and a heavy inert gas insulated layer formed in the walls of the blast tube itself to provide the additional thermal insulation properties to protect any non-propulsive sub-systems positioned in the void space around the blast tube. A void space in the walls of the blast tube contains an inert gas such as Argon, Krypton, Xenon or a synthetic inert gas having a density of at least 1.5 kg/m.sup.3 and a thermal conductivity Tcond_gas of no greater than two-thirds the thermal conductivity of air Tcond_air to form the heavy inert gas insulation layer.

An insulated blast tube includes an insulating layer of a burn resistant material such as phenolic resin formed on an interior surface of the blast tube to provide the necessary erosion and thermal insulation properties to protect the blast tube and a heavy inert gas insulated layer formed in the walls of the blast tube itself to provide the additional thermal insulation properties to protect any non-propulsive sub-systems positioned in the void space around the blast tube. A void space in the walls of the blast tube contains an inert gas such as Argon, Krypton, Xenon or a synthetic inert gas having a density of at least 1.5 kg/m.sup.3 and a thermal conductivity Tcond_gas of no greater than two-thirds the thermal conductivity of air Tcond_air to form the heavy inert gas insulation layer.

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 (6) for feeding a rocket engine (2) at least with a first liquid propellant, the feed circuit including at least one first heat exchanger (18) suitable for being connected to a cooling circuit (17) for cooling at least one heat source in order to cool said heat source by transferring heat to the first propellant, and, in addition, downstream from said first heat exchanger, a branch passing through a second heat exchanger.

A turbine engine wall having a cold side and a hot side and including a plurality of cooling orifices for enabling air flowing on the cold side of the wall to penetrate to the hot side at least some of the cooling orifices being plugged by a plugging material so as to define a minimum level of porosity for the wall corresponding to putting the turbine engine into service, and the plugged cooling orifices being suitable for being unplugged progressively throughout the lifetime of the turbine engine in order to define a maximum level of porosity for the wall corresponding to an end of lifetime for the turbine engine, the plugging being performed by alternating at least one of the following rows or lines: circumferential rows, axial rows, diagonal lines, so as to lie in the range one-third to one-half of the maximum porosity.

A turbine engine wall having a cold side and a hot side and including a plurality of cooling orifices for enabling air flowing on the cold side of the wall to penetrate to the hot side at least some of the cooling orifices being plugged by a plugging material so as to define a minimum level of porosity for the wall corresponding to putting the turbine engine into service, and the plugged cooling orifices being suitable for being unplugged progressively throughout the lifetime of the turbine engine in order to define a maximum level of porosity for the wall corresponding to an end of lifetime for the turbine engine, the plugging being performed by alternating at least one of the following rows or lines: circumferential rows, axial rows, diagonal lines, so as to lie in the range one-third to one-half of the maximum porosity.

A cooling mechanism includes a bottom wall (22) in contact with a combustion chamber, an upper wall (30), and a cooling passage (40) arranged between the bottom wall (22) and the upper wall (30). The cooling passage (40) includes a first passage (50) extending to a first direction, a second passage (60) extending to the first direction, and a connection section (70) connected with the first passage (50) and the second passage (60). The second passage (60) is arranged to have an offset to the first passage (50) in a second direction perpendicular to the first direction and along the bottom wall (22).