US5267437A — Dual mode rocket engine

Claim 1

Independent Claim 354 words

A rocket engine comprising: a combustion chamber/nozzle assembly having a combustion chamber, and a nozzle, said nozzle integral with said combustion chamber; a plurality of tanks containing cryogenic propellants at a pressure greater than ambient pressure; a plurality of first supply lines equal to the number of tanks to pump propellant from the tanks, one of said first supply lines connected to each of said tanks, and each first supply line having a propellant shutoff valve therein; a plurality of primary pumps equal to the number of tanks, each primary pump having a primary pump inlet and a primary pump outlet, each of the primary pump inlets connected to one of said first supply lines; primary turbine means for driving the primary pumps with one of the propellants, said primary turbine means having a primary turbine inlet and a primary turbine outlet; a primary turbine discharge line connecting the primary turbine outlet to the combustion chamber; a boost pump turbine supply line connected to the primary turbine inlet; a first connecting line; a second connecting line connecting the primary pump outlet of one of the primary pumps to the primary turbine inlet; a boost pump turbine discharge line connected to the primary turbine discharge line; a plurality of boost pumps equal to the number of tanks, one of said boost pumps located in each of said tanks, each boost pump including a boost pump turbine having a boost pump turbine inlet and a boost pump turbine outlet, one of said boost pumps having said boost pump turbine supply line connected to the boost pump turbine inlet thereof and the first connecting line connected to the boost pump turbine outlet thereof, and the other boost pump having the first connecting line connected to the boost pump turbine inlet thereof and the boost pump turbine discharge line connected to the boost pump turbine outlet thereof; and, a turbine flow control valve in said primary turbine discharge line between the primary turbine outlet and said boost pump turbine discharge line for selectively directing propellant through the primary turbine discharge line and the boost pump turbine supply line.

Claim 2

Dependent Claim 354 words

A rocket engine comprising: a combustion chamber/nozzle assembly having a combustion chamber, and a nozzle, said nozzle integral with said combustion chamber; a plurality of tanks containing cryogenic propellants at a pressure greater than ambient pressure; a plurality of first supply lines equal to the number of tanks to pump propellant from the tanks, one of said first supply lines connected to each of said tanks, and each first supply line having a propellant shutoff valve therein; a plurality of primary pumps equal to the number of tanks, each primary pump having a primary pump inlet and a primary pump outlet, each of the primary pump inlets connected to one of said first supply lines; primary turbine means for driving the primary pumps with one of the propellants, said primary turbine means having a primary turbine inlet and a primary turbine outlet; a primary turbine discharge line connecting the primary turbine outlet to the combustion chamber; a boost pump turbine supply line connected to the primary turbine inlet; a first connecting line; a second connecting line connecting the primary pump outlet of one of the primary pumps to the primary turbine inlet; a boost pump turbine discharge line connected to the primary turbine discharge line; a plurality of boost pumps equal to the number of tanks, one of said boost pumps located in each of said tanks, each boost pump including a boost pump turbine having a boost pump turbine inlet and a boost pump turbine outlet, one of said boost pumps having said boost pump turbine supply line connected to the boost pump turbine inlet thereof and the first connecting line connected to the boost pump turbine outlet thereof, and the other boost pump having the first connecting line connected to the boost pump turbine inlet thereof and the boost pump turbine discharge line connected to the boost pump turbine outlet thereof; and, a turbine flow control valve in said primary turbine discharge line between the primary turbine outlet and said boost pump turbine discharge line for selectively directing propellant through the primary turbine discharge line and the boost pump turbine supply line.

Claim 3

Dependent Claim 42 words

The rocket engine of claim 1 wherein the boost pump turbine discharge line includes means for preventing propellant from flowing from the primary turbine discharge line to any of said boost pump turbine outlets directly through the boost pump turbine discharge line.

Claim 4

Dependent Claim 57 words

The rocket engine of claim 2 further comprising a bypass line connecting the primary turbine inlet directly to said primary turbine discharge line, said bypass line connected to said primary turbine discharge line between said primary turbine outlet and said turbine flow control valve, said bypass line including means for selectively controlling the flow of propellant therethrough.

Claim 5

Dependent Claim 52 words

The rocket engine of claim 3 further comprising a pressurization line connecting one of the tanks to the boost pump turbine discharge line, said pressurization line including a check valve for preventing propellant from flowing from one of said tanks to the boost pump turbine discharge line directly through the pressurization line.

Claim 6

Dependent Claim 320 words

In a rocket engine having a combustion chamber, a nozzle integral with said combustion chamber, a first tank containing a first cryogenic propellant and a first boost pump, a second tank containing a second cryogenic propellant and a second boost pump, each propellant at a tank pressure in excess of ambient pressure, and a plurality of primary pumps, each primary pump having an inlet for receiving one of the propellants and an outlet for discharging one of the propellants, a method of efficiently cooling the primary pumps to desired temperatures near the temperatures of the propellants in the first and second tanks upon starting the rocket engine, said method comprising: exposing the inlets of the primary pumps to said tank pressures and said outlets of the primary pumps to ambient pressure to create a pressure differential across the inlet and outlet of each of the primary pumps, thereby initiating flow of the propellants through the primary pumps to achieve cooling thereof; discharging the propellants from the primary pumps; vaporizing the first propellant after discharge from one of said primary pumps; using the vaporized first propellant to drive the boost pumps to increase the pressure differential across the primary pumps; mixing the second propellant flowing from one of the primary pumps with the first propellant flowing from the boost pumps and combusting the mixed propellants in the combustion chamber to produce heat and thrust; flowing the first propellant discharging from one of the primary pumps into close proximity with the combustion chamber and nozzle to heat the first propellant discharged from one of the primary pump outlets and increase the vaporization rate thereof to provide additional energy to the vaporized first propellant to drive the boost pumps; maintaining the flow of propellants through the primary pumps until the primary pumps have been cooled to the desired temperature by the propellants; and using the heated first propellant to drive the primary pumps.

Claim 7

Dependent Claim 320 words

In a rocket engine having a combustion chamber, a nozzle integral with said combustion chamber, a first tank containing a first cryogenic propellant and a first boost pump, a second tank containing a second cryogenic propellant and a second boost pump, each propellant at a tank pressure in excess of ambient pressure, and a plurality of primary pumps, each primary pump having an inlet for receiving one of the propellants and an outlet for discharging one of the propellants, a method of efficiently cooling the primary pumps to desired temperatures near the temperatures of the propellants in the first and second tanks upon starting the rocket engine, said method comprising: exposing the inlets of the primary pumps to said tank pressures and said outlets of the primary pumps to ambient pressure to create a pressure differential across the inlet and outlet of each of the primary pumps, thereby initiating flow of the propellants through the primary pumps to achieve cooling thereof; discharging the propellants from the primary pumps; vaporizing the first propellant after discharge from one of said primary pumps; using the vaporized first propellant to drive the boost pumps to increase the pressure differential across the primary pumps; mixing the second propellant flowing from one of the primary pumps with the first propellant flowing from the boost pumps and combusting the mixed propellants in the combustion chamber to produce heat and thrust; flowing the first propellant discharging from one of the primary pumps into close proximity with the combustion chamber and nozzle to heat the first propellant discharged from one of the primary pump outlets and increase the vaporization rate thereof to provide additional energy to the vaporized first propellant to drive the boost pumps; maintaining the flow of propellants through the primary pumps until the primary pumps have been cooled to the desired temperature by the propellants; and using the heated first propellant to drive the primary pumps.

Annuity Stage	Window Opens	Payment Deadline	Surcharge Period Until	Estimated Fee	Status
3.5 Years (Year 4)	1996-12-06	1997-06-07	1997-12-07	$2000.00	Window Closed / Lapsed
7.5 Years (Year 8)	2000-12-06	2001-06-07	2001-12-07	$3760.00	Window Closed / Lapsed
11.5 Years (Year 12)	2004-12-06	2005-06-07	2005-12-07	$7700.00	Window Closed / Lapsed

Dual mode rocket engine

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