Spacecraft propellant tank and its application method

Abstract

The invention concerns a spacecraft propellant tank designed for storing and supplying fuel to the propulsion system, as well as its application method, which is related to space technology and particularly to the design of tanks for storing liquid propellants within rocket engines. The tank in question comprises a body featuring at least one convex edge that creates a corner capillary and an intake port positioned on one of these convex edges. Additionally, the tank's interior surface is treated to be hydrophilic. The objective of this invention is to develop a spacecraft propellant tank characterized by its lightweight and straightforward design. It facilitates the creation of various propellant tank shapes while minimizing the quantity of internal components that reduce useful volume. It also ensures consistent and reliable functionality of the propellant system with continuous supply of propellant to the intake port, even in the absence of gravity and during multiple activations of the propulsion system. The claimed invention achieves its objective by implementing technical results that include enhancing the efficiency of propellant supply within the tank to the intake port, improving reliability, augmenting the useful internal volume, and offering a variety of tank shapes. The stated objectives are met in particular through the hydrophilic properties of the propellant tank's inner surface, the design of the propellant tank shape featuring at least one convex edge that forms a corner capillary, and positioning of the intake port on one of the convex edges that forms the corner capillary. The application method for the spacecraft propellant tank involves initially filling the tank with liquid propellant and pressurizing gas through the intake port prior to spacecraft launch. Subsequently, once the spacecraft is in orbit, propellant is conveyed from the tank to the propulsion system through the intake port. The movement of liquid propellant towards the intake port is facilitated by the pressure of the pressurizing gas and a localized pressure reduction generated during propellant discharge. This reduction occurs in the corner capillary near the intake port relative to the pressure of the propellant in other sections of at least one corner capillary, thus establishing a flow of propellant through the corner capillaries towards the intake port area.

Dependencies are inferred from the claim text.