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;
an oxidizer throttle receiving the oxidizer from the oxidizer supply line, the oxidizer supply line splitting into a first partial oxidizer supply line, a second partial oxidizer supply line, and a third partial oxidizer supply line;
a turbine mounted to a turbine shaft along a longitudinal centerline of the LRE at a longitudinal proximal location of the LRE, the turbine rotating with the turbine shaft about the longitudinal centerline;
a gas duct in fluid communication with the turbine;
a main combustion chamber having a main chamber injector head in fluid communication with both the gas duct and the second partial oxidizer supply line;
a nozzle positioned longitudinally distal to the main combustion chamber and in fluid communication with the main combustion chamber;
a preburner combustion chamber positioned axially distal to the main combustion chamber and having a preburner injector head in fluid communication with both a first cooling manifold and a second cooling manifold, wherein the preburner combustion chamber forms an annulus around the main combustion chamber;
the first cooling manifold in fluid communication with the fuel supply line;
the second cooling manifold positioned axially between the preburner combustion chamber and the main combustion chamber, the second cooling manifold in fluid communication with the first partial oxidizer supply line; and
a third cooling manifold in fluid communication with the third partial oxidizer supply line and a nozzle throat of the nozzle;
wherein:
the fuel received by the first cooling manifold from the fuel supply line receives thermal energy from the nozzle and is supplied to the preburner injector head;
the oxidizer received by the second cooling manifold from the first partial oxidizer supply line receives thermal energy from at least one of the gas duct, the main combustion chamber, and the preburner combustion chamber, the oxidizer supplied to the preburner injector head;
the preburner injector head injects the fuel with the oxidizer to create preburner combustion exhaust products which flow from the preburner combustion chamber to the turbine and drive the turbine about the turbine shaft, the preburner combustion exhaust products flowing from the turbine to the gas duct; and
the main combustion chamber injector head injects the preburner combustion exhaust products received from the gas duct with oxidizer received from the second partial oxidizer supply line to produce an LRE thrust directed through the nozzle exit.
2. The liquid rocket engine of claim 1 , wherein the fuel is methane.
3. The liquid rocket engine of claim 1 , wherein the main combustion chamber comprises a main chamber ignitor which ignites the preburner combustion exhaust products with the oxidizer.
4. The liquid rocket engine of claim 1 , wherein the turbine is a centrifugal turbine.
5. The liquid rocket engine of claim 1 , further comprising a set of turbine guide vanes operating to direct the preburner combustion exhaust products into the turbine.
6. The liquid rocket engine of claim 1 , wherein the gas duct is positioned axially between the turbine and the main combustion chamber.
7. 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.
8. 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.
9. A liquid rocket engine (LRE) comprising:
a fuel supply line containing a fuel;
an oxidizer supply line containing an oxidizer;
an oxidizer throttle receiving the oxidizer from the oxidizer supply line, 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 at a longitudinal proximal location of the LRE, the turbine rotating with the turbine shaft about the longitudinal centerline;
a main combustion chamber having an injector head in fluid communication with the turbine and at least one of the plurality of partial oxidizer supply lines;
a nozzle positioned longitudinally distal to the main combustion chamber and in fluid communication with the main combustion chamber; and
a preburner combustion chamber positioned axially distal to the main combustion chamber and having a preburner injector head, the preburner combustion chamber forming an annulus around the main combustion chamber;
wherein:
the preburner injector head injects a gasified fuel with a gasified oxidizer to create preburner combustion exhaust products which flow from the preburner combustion chamber to the turbine and drive the turbine about the turbine shaft, the preburner combustion exhaust products flowing from the turbine to the main combustion chamber; and
the main combustion chamber injector head injects the preburner combustion exhaust products received from the turbine with oxidizer to produce an LRE thrust directed through the nozzle exit.
10. The liquid rocket engine of claim 9 , further comprising a gas duct positioned axially between the turbine and the main combustion chamber.
11. The liquid rocket engine of claim 10 , 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.
12. The liquid rocket engine of claim 11 , 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.
13. The liquid rocket engine of claim 9 , wherein the fuel is methane.
14. The liquid rocket engine of claim 9 , wherein the turbine is a centrifugal turbine.
15. The liquid rocket engine of claim 9 , further comprising a set of turbine guide vanes operating to direct the preburner combustion exhaust products into the turbine.
16. 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 at a longitudinal proximal location of the LRE;
a main combustion chamber having an injector head in fluid communication with the turbine and at least one of the plurality of partial oxidizer supply lines;
a nozzle positioned longitudinally distal to the main combustion chamber and in fluid communication with the main combustion chamber; and
a preburner combustion chamber positioned axially distal to the main combustion chamber and having a preburner injector head, the preburner combustion chamber forming an annulus around the main combustion chamber;
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, a gasified fuel with a gasified oxidizer to create preburner combustion exhaust products which flow from the preburner combustion chamber to the turbine;
rotating the turbine about the turbine shaft;
flowing 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 directed through the nozzle exit.
17. The method of claim 16 , the LRE further comprising a gas duct positioned axially between the turbine and the main combustion chamber, 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 method 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 method of claim 16 , wherein:
the turbine is a centrifugal turbine; and
the fuel is methane.
20. The method of claim 16 , the LRE further comprising a set of turbine guide vanes operating to direct the preburner combustion exhaust products into the turbine.