Staged combustion liquid rocket engine cycle with the turbopump unit and preburner integrated into the structure of the combustion chamber
Devices and methods of rocket propulsion are disclosed. In one aspect, a staged combustion liquid rocket engine with preburner and turbopump unit (TPU) integrated into the structure of the combustion chamber is described. An initial propellant mixture is combusted in a preburner combustion chamber formed as an annulus around a main combustion chamber, the combustion products from the preburner driving the turbine of the TPU and subsequently injected into the main combustion chamber for secondary combustion along with additional propellants, generating thrust through a supersonic nozzle. The preburner inner cylindrical wall is shared with the outer cylindrical wall of the engine's main combustion chamber and the turbine is axially aligned with the main combustion chamber. Liquid propellants supplied to the engine are utilized for regenerative cooling of the combustion chamber and preburner, where the liquid propellants are gasified in cooling manifolds before injection into the preburner and main combustion chamber.
1. A liquid rocket engine (LRE) comprising:
a fuel supply line containing a fuel;
an oxidizer supply line containing an oxidizer, the oxidizer supply line splitting into a plurality of partial oxidizer supply lines;
a turbine mounted to a turbine shaft along a longitudinal centerline of the LRE and rotating with the turbine shaft about the longitudinal centerline;
a main combustion chamber having a main combustion chamber injector head in fluid communication with the turbine and in fluid communication with at least one of the plurality of partial oxidizer supply lines;
a flow collimator comprising a set of apertures;
a nozzle in fluid communication with the main combustion chamber; and
a preburner combustion chamber having a preburner injector head;
wherein:
the preburner injector head injects the fuel with the oxidizer to create preburner combustion exhaust products which drive the turbine about the turbine shaft, the preburner combustion exhaust products flowing from the turbine to the main combustion chamber via a gas duct;
the flow collimator is in fluid communication with the gas duct and the main combustion chamber; and
the main combustion chamber injector head injects the preburner combustion exhaust products received from the turbine with the oxidizer to produce an LRE thrust.
2. The liquid rocket engine of claim 1 , wherein the gas duct is positioned axially between the turbine and the main combustion chamber.
3. The liquid rocket engine of claim 1 , wherein at least some of the oxidizer becomes a gasified oxidizer due to thermal energy transfer from at least one of the gas duct, the main combustion chamber, and the preburner combustion chamber.
4. The liquid rocket engine of claim 1 , wherein at least some of the fuel becomes a gasified fuel due to thermal energy transfer from the nozzle to the at least some of the fuel, the gasified fuel supplied to the preburner injector head.
5. The liquid rocket engine of claim 1 , wherein the flow collimator is positioned longitudinally proximal to the main combustion chamber injector head and positioned longitudinally distal to the gas duct.
6. The liquid rocket engine of claim 5 , further comprising a lattice structure positioned proximal to the main combustion chamber injector head.
7. The liquid rocket engine of claim 1 , further comprising a set of supplemental fuel injection ports positioned distal to the main combustion chamber infector head.
8. The liquid rocket engine of claim 1 , wherein:
the turbine is a centrifugal turbine;
the preburner combustion chamber forms an annulus around the main combustion chamber; and
the fuel is methane.
9. A method of operating a liquid rocket engine (LRE) comprising:
providing a LRE comprising:
a fuel supply line containing a fuel;
an oxidizer supply line containing an oxidizer;
an oxidizer throttle;
a turbine mounted to a turbine shaft along a longitudinal centerline of the LRE;
a main combustion chamber having a main combustion chamber injector head in fluid communication with each of the turbine and at least one of the plurality of partial oxidizer supply lines;
a gas duct;
a flow collimator comprising a set of apertures, the flow collimator in fluid communication with the gas duct and with the main combustion chamber;
a nozzle in fluid communication with the main combustion chamber; and
a preburner combustion chamber having a preburner injector head;
receiving the oxidizer from the oxidizer supply line;
splitting the oxidizer supply line into a plurality of partial oxidizer supply lines;
injecting, using the preburner injector head, the fuel with the oxidizer to create preburner combustion exhaust;
rotating the turbine about the turbine shaft;
flowing, via the gas duct, the preburner combustion exhaust products from the turbine to the main combustion chamber; and
injecting, using the main combustion chamber injector head, the preburner combustion exhaust products received from the turbine with oxidizer to produce an LRE thrust.
10. The method of claim 9 , wherein:
the gas duct is positioned axially between the turbine and the main combustion chamber; and
at least some of the oxidizer becomes a gasified oxidizer due to thermal energy transfer from at least one of the gas duct, the main combustion chamber, and the preburner combustion chamber.
11. The method of claim 9 , wherein at least some of the fuel becomes a gasified fuel due to thermal energy transfer from the nozzle to the at least some of the fuel, the gasified fuel supplied to the preburner injector head.
12. The method of claim 9 , wherein the flow collimator is positioned longitudinally proximal to the main combustion chamber injector head and positioned longitudinally distal to the gas duct.
13. The method of claim 12 , the LRE further comprising a lattice structure positioned proximal to the main combustion chamber injector head.
14. The method of claim 9 , the LRE further comprising a set of supplemental fuel injection ports positioned distal to the main combustion chamber injector head.
15. The method of claim 9 , wherein:
the turbine of the LRE is a centrifugal turbine;
the preburner combustion chamber forms an annulus around the main combustion chamber; and
the fuel is methane.
16. A rocket engine comprising:
a fuel supply line configured to supply a fuel;
an oxidizer supply line configured to supply an oxidizer, the oxidizer supply line comprising a set of partial oxidizer supply lines;
a turbine mounted to a turbine shaft along a longitudinal centerline of the rocket engine, the turbine configured to rotate with the turbine shaft about the longitudinal centerline;
a main combustion chamber comprising a main combustion chamber injector head, the main combustion chamber in fluid communication with the turbine and in fluid communication with at least one of the set of partial oxidizer supply lines;
a flow collimator positioned longitudinally proximal to the main combustion chamber injector head, the flow collimator comprising a set of apertures positioned longitudinally distal to the gas duct, the flow collimator in fluid communication with a gas duct and with the main combustion chamber; and
a preburner combustion chamber comprising a preburner injector head;
wherein:
the preburner injector head injects the fuel with the oxidizer to create preburner combustion exhaust products which drive the turbine about the turbine shaft; and
the main combustion chamber injector head injects the preburner combustion exhaust products received from the turbine with the oxidizer to produce a rocket engine thrust.
17. The rocket engine of claim 16 , wherein at least some of the oxidizer becomes a gasified oxidizer due to thermal energy transfer from at least one of the gas duct, the main combustion chamber, and the preburner combustion chamber.
18. The rocket engine of claim 16 , wherein at least some of the fuel becomes a gasified fuel due to thermal energy transfer from the nozzle to the at least some of the fuel, the gasified fuel supplied to the preburner injector head.
19. The rocket engine of claim 16 , further comprising:
a lattice structure positioned proximal to the main combustion chamber injector head;
a nozzle in fluid communication with the main combustion chamber; and
a set of supplemental fuel injection ports positioned distal to the main combustion chamber injector head.
20. The rocket engine of claim 16 , wherein the preburner combustion chamber forms an annulus around the main combustion chamber.