Hybrid-cycle liquid propellant rocket engine
View Patent ↗Systems and methods are described herein for a hybrid liquid propellant rocket engine. In an embodiment, the engine includes a first pump powered by a first turbine, a second pump powered by a second turbine, and a gas generator. An output of the gas generator is connected to the first turbine and the second turbine. The engine further includes a third pump powered by a third turbine, a fourth pump powered by a fourth turbine, and a nozzle having an expander cycle in a wall and a combustion chamber. An output of the third pump is connected to the expander cycle and an output of the wall is connected to the third turbine and the fourth turbine. An output of the fourth pump, an output of the third turbine, and an output of the fourth turbine are connected to the combustion chamber.
1. A hybrid rocket engine comprising:
a first pump powered by a first turbine;
a second pump powered by a second turbine;
a gas generator connected to the first pump and the second pump, wherein an output of the gas generator is connected to the first turbine and the second turbine to provide mechanical power to the first pump and the second pump;
a third pump powered by a third turbine;
a fourth pump powered by a fourth turbine; and
a nozzle comprising an expander cycle in a wall and a combustion chamber;
wherein an output of the third pump is connected to the expander cycle, and an output of the expander cycle is connected to the third turbine and the fourth turbine to provide mechanical power to the third pump and the fourth pump, and
wherein an output of the fourth pump, an output of the third turbine, and an output of the fourth turbine are connected to the combustion chamber.
2. The hybrid rocket engine of claim 1 , wherein an output of the first turbine and an output of the second turbine are expelled as exhaust.
3. The hybrid rocket engine of claim 1 , wherein the first turbine, the second turbine, the third turbine, and the fourth turbine are isolated from one another.
4. The hybrid rocket engine of claim 1 , wherein the first pump and the third pump are configured to pump a fuel, and the second pump and the fourth pump are configured to pump an oxidizer.
5. The hybrid rocket engine of claim 1 , wherein the gas generator is configured to combust a portion of a fuel from the first pump and a portion of an oxidizer from the second pump to produce a gas that powers the first pump and the second pump.
6. A hybrid rocket engine comprising:
a first pump powered by a first turbine and a second turbine, wherein the first turbine and the second turbine are mechanically connected;
a second pump powered by a third turbine and a fourth turbine, wherein the third turbine and the fourth turbine are mechanically connected;
a gas generator connected to the first pump and the second pump, wherein an output of the gas generator is connected to the first turbine and the third turbine to provide mechanical power to the first pump and the second pump; and
a nozzle comprising an expander cycle in a wall and a combustion chamber;
wherein an output of the first pump is connected to the expander cycle, and an output of the expander cycle is connected to the second turbine and the fourth turbine to provide mechanical power to the first pump and the second pump, and
wherein an output of the second pump, an output of the third turbine, and an output of the fourth turbine are connected to the combustion chamber.
7. The hybrid rocket engine of claim 6 , wherein an output of the first turbine and an output of the second turbine are expelled as exhaust.
8. The hybrid rocket engine of claim 6 , wherein the first turbine and the second turbine are connected in series to the first pump, and the third turbine and the fourth turbine are connected in series to the second pump.
9. The hybrid rocket engine of claim 6 , wherein the first pump is configured to pump a fuel and the second pump is configured to pump an oxidizer.
10. The hybrid rocket engine of claim 6 , wherein the gas generator is configured to combust a portion of a fuel from the first pump and a portion of an oxidizer from the second pump to produce a gas that powers the first pump and the second pump.
11. A method of combusting a fuel and an oxidizer in a hybrid rocket engine, the method comprising:
pumping, at a first pump, a first portion of a fuel to a gas generator;
pumping, at a second pump, a first portion of an oxidizer to a gas generator;
combusting, at a gas generator, the first portion of the fuel and the first portion of the oxidizer to power the first pump and the second pump, wherein the first pump is mechanically connected to a first turbine and the second pump is mechanically connected to a second turbine;
pumping, at the first pump, a second portion of the fuel to a third pump;
pumping, at the second pump, a second portion of the oxidizer to a fourth pump;
pumping, at the third pump, the second portion of the fuel to an expander cycle in a wall of a nozzle and into a third turbine and a fourth turbine, wherein an output of the third turbine and an output of the fourth turbine are provided to a combustion chamber in the nozzle, and wherein the third turbine is mechanically connected to the third pump and the fourth turbine is mechanically connected to the fourth pump;
pumping, at the fourth pump, the second portion of the oxidizer into the combustion chamber;
combusting, at the combustion chamber, the second portion of the fuel and the second portion of the oxidizer to provide thrust from the nozzle.
12. The method of claim 11 , further comprising expelling an output of the first turbine and an output of the second turbine as exhaust.
13. The method of claim 11 , wherein the expander cycle changes a phase of the fuel.
14. The method of claim 11 , wherein combusting the first portion of the fuel and the first portion of the oxidizer generates a gas that powers the first pump and the second pump.
15. The method of claim 11 , further comprising closing a first valve to block the passage of fuel to the gas generator, closing a second valve to block the passage of oxidizer to the gas generator, and using the hybrid rocket engine as an expander cycle rocket engine.
16. A method of combusting a fuel and an oxidizer in a hybrid rocket engine, the method comprising:
pumping, at a first pump, a first portion of a fuel to a gas generator, wherein the first pump is mechanically connected to a first turbine and a second turbine;
pumping, at a second pump, a first portion of an oxidizer to a gas generator, wherein the second pump is mechanically connected to a third turbine and a fourth turbine;
combusting, at a gas generator, the first portion of the fuel and the first portion of the oxidizer to power the first pump and the second pump;
pumping, at the first pump, a second portion of the fuel to an expander cycle in a wall of a nozzle and into the second turbine and the fourth turbine, wherein an output of the second turbine and an output of the fourth turbine are provided to a combustion chamber in the nozzle;
pumping, at the second pump, the second portion of the oxidizer into the combustion chamber;
combusting, at the combustion chamber, the second portion of the fuel and the second portion of the oxidizer to provide thrust from the nozzle.
17. The method of claim 16 , further comprising expelling an output of the first turbine and an output of the third turbine as exhaust.
18. The method of claim 16 , wherein the expander cycle changes a phase of the fuel.
19. The method of claim 16 , wherein combusting the first portion of the fuel and the first portion of the oxidizer generates a gas that powers the first pump and the second pump.
20. The method of claim 16 , further comprising closing a first valve to block the passage of fuel to the gas generator, closing a second valve to block the passage of oxidizer to the gas generator, and using the hybrid rocket engine as an expander cycle rocket engine.