Power plants – Reaction motor – Turborocket
Reexamination Certificate
2000-11-01
2004-02-17
Yu, Justine R. (Department: 3746)
Power plants
Reaction motor
Turborocket
C060S039465, C060S257000, C060S258000, C060S259000
Reexamination Certificate
active
06691504
ABSTRACT:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT
Not Applicable
REFERENCE TO MICROFICHE APPENDIX
Not Applicable
BACKGROUND OF THE INVENTION
The present invention is a gaseous-fuel breathing rocket engine power plant for accelerating vehicles such as aircraft and spacecraft by reaction thrusting. The principal use of the present invention is on aircraft and spacecraft having a large reservoir of gaseous fuel, which may be combustible by oxidation or in some other exothermic reaction.
The preferred gaseous fuel for the engine is a gaseous fuel that contains hydrogen gas. The reservoir containing such gaseous fuel may be the gas retaining structures of an airship, such as gas bags, wherein the gaseous fuel serves as the lifting gas.
The types of propulsion systems which create a propulsion force known as thrust to propel vehicles at high altitudes are the rocket motor and the jet engine. The propulsion force is the reaction force arising from increasing the backward momentum of a mass by the action of the propulsion system. In the case of the rocket motor, the rearward ejected mass comes from the propellant chemicals carried with the vehicle, and the backward momentum from the reaction between those propellant chemicals. In the case of the jet engine, addition of heat energy to a controlled flow of air passing through the jet engine increases the backward momentum of the airflow.
Some of the features of the present invention disclosed here relate to features of both jet engines and rocket motors. Unlike conventional jet engines which intake air, the present invention intakes gaseous fuel. If the gaseous fuel intaken does not have sufficient density for efficient combustion, it is compressed within the engine to achieve such combustion. Also, unlike conventional jet engines, the combustion of the gaseous fuel compressed by the turbine compressors takes place with a stored oxidizer which is injected into the gaseous fuel stream.
The use of gaseous hydrogen as fuel for power plants which compress air with turbine compressors is known from U.S. Pat. No. 5,012,640, The Combined Air-Hydrogen Turbo-Rocket Power Plant. The power plant disclosed in that patent, however, uses evaporating liquid hydrogen to drive a turbine which powers a turbine compressor to compress air into which gaseous hydrogen is injected for combustion, and does not use a turbine compressor to compress gaseous fuel or use liquid oxygen to drive the turbine. Also, that power plant does not use stored oxidizer to burn the hydrogen, but uses the air which has been compressed for such combustion. A power plant similar to the Combined Air-Hydrogen Turbo-Rocket Power Plant disclosed in U.S. Pat. No. 5,012,640 is disclosed in French Patent No. 2,215,538, but the principal difference between them appears to be the type and placement of the turbine which drives the air compressor. The drawings of U.S. Pat. No. 5,012,640 and French Patent No. 2,215,583 provide the needed current art with respect to driving turbine compressors with evaporating liquid gases, and are used for that purpose in this application.
A gaseous-fuel breathing turbo-rocket thruster has previously been disclosed by the present applicant in United States Patent Application No. 09/321,796, presently allowed and awaiting issue and publication, and International Application No. PCT/US00/09617, recently transmitted to the International Bureau. However, the thruster disclosed in Application No. 09/321,796 uses exhaust gases generated by the combustion of the intaken gaseous-fuel to drive a coaxial turbine which in turn drives the gaseous-fuel compressor. Such utilization of the energetic exhaust of said thruster may needlessly diminish the exit velocity of such exhaust.
The present invention has elements that are covered generally by Current United States Class 60, power plants, particularly subclass 246, and International Class: F02M 067/00; F02M 002/08; F01D 005/20.
BRIEF SUMMARY OF THE INVENTION
A combined gaseous-fuel and oxygen rocket engine is disclosed in which the oxygen driven turbine or electric motor drives the rotor wheel blades of the axial gaseous-fuel compressor stages. The rotor stages are located downstream of a stator vane structure and are driven by gaseous oxygen passing across the turbine blades. The oxygen is subsequently injected into an gaseous-fuel duct surrounding the axial gaseous-fuel compressor and defining an gaseous-fuel flow path having an gaseous-fuel inlet. The oxygen and gaseous-fuel mixture is ignited and the burned gases are expanded through a converging-diverging exhaust nozzle.
In the case of the gaseous-fuel compressor being powered by an oxygen driven turbine, the oxygen is supplied to the turbine from a liquid oxygen reservoir via at least one oxygen pump with the liquid passing through a heat exchanger to raise the temperature of the oxygen, thereby causing it to vaporize. The gaseous oxygen then passes to and drives a turbine, which in turn drives the axial gaseous-fuel compressor.
The oxygen pump may be driven electrically, or by an auxiliary turbine, again powered by gaseous oxygen, or may be mounted in the hub of the axial gaseous-fuel compressor and be driven directly by the compressor rotor wheel.
In the power plant according to the invention, each compressor rotor wheel may be driven electrically, or by at least one axial flow turbine rotor stage located outside the compressed gaseous-fuel duct in an annular chamber surrounding the duct, or by a turbine which is coaxially located with respect to the gaseous-fuel compressor and is connected to the compressor by a generally axially extending shaft.
In alternative embodiments of the invention, more than one turbine rotor blade may be associated with each of the axial compressor rotor blades and the axial compressor may comprise more than one rotor stage. If a plurality of compressor rotor stages are utilized, adjacent stages may rotate in the same direction, or they may rotate in opposite directions depending upon the orientation of the turbine rotor blades.
The multiple stages of the axial compressor may be located in a common annular chamber, or they may be located in separate annular chambers which may be connected to the oxygen supply system either in parallel or in series.
REFERENCES:
patent: 4224790 (1980-09-01), Christensen
patent: 5012640 (1991-05-01), Mirville
patent: 5014507 (1991-05-01), Rice
patent: 5063734 (1991-11-01), Morris
patent: 5101622 (1992-04-01), Bond
patent: 5778658 (1998-07-01), Lamando
patent: 6148609 (2000-11-01), Provitola
Rodriguez William H.
Yu Justine R.
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