Systems and methods for design, manufacture, and test of space thruster propulsion systems
The disclosed methods, systems, and kits provide the ability to deliver entire clean sheet designs from concept to first hot fire in under six weeks with instant specific impulses above 330 seconds in some of our engines. In examples, thrusters can be delivered that are at less than half of the mass budget allowable for them and they can be delivered in weeks.
1 . A method of manufacturing a rocket, the method comprising: additively manufacturing a combustion chamber;
additively manufacturing an injector that is integrally coupled with the combustion chamber; and
additively manufacturing propellant management structures that are integrally coupled with the injector and the combustion chamber that include additively manufacturing a regenerative jacket around the combustion chamber that includes:
additively manufacturing a plurality of ribs that couple with a hot wall of the combustion chamber and a cold wall of the combustion chamber and are oriented to set a propellant flow to impinge upon a plurality of fins; and
additively manufacturing the plurality of fins that extend from the hot wall of the combustion chamber towards the cold wall of the combustion chamber and protrude into one or more flow paths between the plurality of ribs, wherein the plurality of ribs and the plurality of fins form a plurality of stages that include:
a first stage from the plurality of stages that includes a first plurality of ribs from the plurality of ribs and a first plurality of fins from the plurality of fins; and
a second stage from the plurality of stages that includes a second plurality of ribs from the plurality of ribs and a second plurality of fins from the plurality of fins, wherein the second stage follows discontinuously from the first stage in the one or more flow paths between the plurality of ribs.
2 . The method of claim 1 , further comprising additively manufacturing a nozzle that is integrally coupled with the combustion chamber.
3 . The method of claim 1 , further comprising additively manufacturing a catalyst that is integrally coupled with a wall of the combustion chamber.
4 . The method of claim 3 , wherein additively manufacturing the catalyst includes additively manufacturing the catalyst based on maximizing a surface area of the catalyst based on a substantially smooth wall and a feature resolution of a machine used to additively manufacture the catalyst.
5 . The method of claim 1 , wherein additively manufacturing at least one of the combustion chamber, the injector, and the propellant management structures include additively manufacturing using a nickel alloy.
6 . The method of claim 1 , wherein additively manufacturing at least one of the combustion chamber, the injector, and the propellant management include additively manufacturing using niobium C103.
7 . The method of claim 1 , wherein additively manufacturing at least one of the combustion chamber, the injector, and the propellant management structures include additively manufacturing using a cobalt chrome alloy.
8 . The method of claim 1 , wherein additively manufacturing at least one of the combustion chamber, the injector, and the propellant management structures include additively manufacturing the combustion chamber using a molybdenum alloy.
9 . The method of claim 1 , wherein additively manufacturing at least one of the combustion chamber, the injector, and the propellant management structures include additively manufacturing the combustion chamber using a rhenium alloy, the method further comprising lining the combustion chamber with at least one layer of iridium.
10 . The method of claim 1 , further comprising additively manufacturing seals between components of the rocket.
11 . The method of claim 10 , further comprising additively manufacturing seal glands for the seals.
12 . The method of claim 10 , further comprising additively manufacturing a hot-gas metal seal between the combustion chamber and injector.
13 . The method of claim 1 , wherein additively manufacturing at least one of the combustion chamber, the injector, and the propellant management structures include forming fine features having dimensions configured to cause the propellant to transition from a liquid phase to a gas phase prior to combustion to promote propellant mixing under a predetermined standard operation of the rocket.
14 . The method of claim 13 , further comprising forming fine features for a rocket fuel and forming fine features for an oxidizer, wherein forming fine features for the rocket fuel is based on a phase transition requirement of the rocket fuel and forming fine features for the oxidizer is based on a phase transition requirement of the oxidizer.
15 . The method of claim 13 , wherein forming fine features includes forming injector elements having an aperture dimension of less than one-hundredth of an inch to improve propellant atomization, reduce local bed loading, restrict quenching in an induction zone and limit hot spots between the injector elements to lower injector face maximum temperatures.
16 . The method of claim 1 , wherein additively manufacturing the propellant management structures includes forming propellant tubing in a honeycomb shape disposed proximate the combustion chamber to transfer heat from the combustion chamber to the propellant during propellant flow in a predetermined standard operation of the rocket.
17 . The method of claim 1 , wherein additively manufacturing the propellant management structures includes forming propellant tubing using materials selected to have a thermal effusivity that heats the propellant and cools the combustion chamber to promote transition of the propellant to a gas phase during a predetermined standard operation of the rocket.
18 . The method of claim 1 , further comprising dispersing ceramic particles in a metal matrix for high temperature strength while additively manufacturing.
19 . The method of claim 1 , wherein additively manufacturing further includes additively manufacturing with refractory metals.
20 . The method of claim 1 , further comprising adjusting a porosity of the rocket based on a proximity to a chamber wall of the combustion chamber.
21 . The method of claim 20 , further comprising additively manufacturing disfavored thermal conduction pathways within at least one of the combustion chamber, the injector, and the propellant management structures,
wherein the disfavored thermal conduction pathways have a thermal resistance that is higher than a thermal resistance of a remainder of the at least one of the combustion chamber, the injector, and the propellant management structures that is not part of the disfavored thermal conduction pathways for directing heat transfer in the rocket.
22 . The method of claim 1 , further comprising additively manufacturing embedded instrumentation that is integrally coupled with the rocket.
23 . The method of claim 1 , wherein additively manufacturing at least one of the combustion chamber, the injector, and the propellant management structures include integrally forming instrument ports configured to receive instrumentation components.
24 . The method of claim 1 , wherein additively manufacturing at least one of the combustion chamber, the injector, and the propellant management structures include integrally forming tube stubs configured to receive at least one of Ground Support Equipment (GSE) components and thruster components.
25 . The method of claim 1 , wherein additively manufacturing the injector includes forming injection elements normal to a face of the injector.
26 . The method of claim 1 , wherein additively manufacturing the injector includes forming baffles with exits normal to a face of the injector, wherein the baffles are configured to dampen acoustics in the rocket.
27 . The method of claim 1 , wherein additively manufacturing the injector includes forming injector elements with integral vorticity generators.
28 . The method of claim 1 ,
wherein additively manufacturing the injector includes forming the injector in a dome shape, and
wherein additively manufacturing the combustion chamber includes forming the combustion chamber in a spherical shape and forming baffles on walls of the combustion chamber.
29 . The method of claim 1 , wherein:
the first plurality of fins from the plurality of fins includes:
a first group of fins from the first plurality of fins oriented parallel to the first plurality of ribs from the plurality of ribs; and
a second group of fins from the first plurality of fins oriented obliquely to the first plurality of ribs from the plurality of ribs; and
the second plurality of fins from the plurality of fins includes:
a first group of fins from the second plurality of fins oriented parallel to the second plurality of ribs from the plurality of ribs; and
a second group of fins from the second plurality of fins oriented obliquely to the second plurality of ribs from the plurality of ribs.
30 . The method of claim 29 , further comprising:
additively manufacturing a plurality of channels between the hot wall of the combustion chamber and the cold wall of the combustion chamber that terminate axially below the plurality of ribs and the plurality of fins and that resume axially above the plurality of ribs and the plurality of fins.