Power plants – Reaction motor – Turborocket
Reexamination Certificate
2001-04-25
2003-09-16
Gartenberg, Ehud (Department: 3746)
Power plants
Reaction motor
Turborocket
Reexamination Certificate
active
06619031
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates to liquid propellant rocket engines and particularly to multi-mode multi-propellant single stage earth to orbit or suborbital rocket engines.
2. Description of the Prior Art
Rocket propulsion for earth to orbit launch vehicles is currently the only practical choice. Known rocket engines, however, operate at efficiencies that are far from the optimum, particularly for single stage earth to orbit operation. While many solutions for the enhancement of liquid fuel rocket engines have been proposed, few of them have been implemented. Most innovations lead to significant design complications and cost increases which offset their potential benefits. As a result, principles discovered early in the 20th century, such as for example multistage rockets, still provide the main basis for modern rocket launchers. It is well known from rocket theory that launcher efficiency can be increased if high thrust, dense propellant, moderate specific impulse engines are used at low altitudes for the initial acceleration, and high specific impulse, lower thrust engines are used for high altitude acceleration and orbiting. Liquid propellant combinations embodying oxidizers and fuels which have been considered are liquid oxygen(LOX)/kerosene and liquid oxygen (LOX)/liquid hydrogen(LH2) engines; LOX/methane and LOX/LH2 engines; and solid propellant motors and LOX/LH2 engines. Rocket engines using such fuel combinations have been built and tested, and most of them are workhorses for modern launch businesses. The use of different pairs of fuels for different hardware units essentially requires a multistage configuration of rocket engines. Single-stage-to-orbit (SSTO) launchers are not feasible because idling groups of engines must be carried. Multistage configurations further result in high launch costs and prevent introduction of reusability. The use a fuel efficient LOX/LH2 engine as the only thruster for SSTO launchers is unlikely to be successful because such an engine is not efficient when used as a low altitude thruster.
U.S. Pat. Nos. 4,771,599 and 4,771,600 disclose tripropellant rocket engines utilizing a tripropellant fuel system in which the propellants are oxygen, hydrogen, and a hydrocarbon. Such an engine will produce the thrust necessary for initial acceleration of significant payloads into low earth orbit. This engine is referred to as a booster or high thrust low altitude engine providing for initial acceleration of the launcher, and it has only one operational mode. It is not advantageous to use this engine for high-speed acceleration and orbiting because it has a lower specific impulse than the conventional LOX/LH2 engine.
U.S. Pat. No. 4,831,818 discloses a dual-fuel, dual-mode single stage rocket engine for earth to orbit operation. The fuels are a high specific impulse fuel such as liquid hydrogen, and a high density-impulse fuel such as liquid methane. Flow of the fuels is said to be controlled by the fuel pumps. The fuels are used to cool the nozzle. The fuels are mixed upstream of the nozzle cooling jacket, and the fuel mixture is fed to the cooling jacket. A method is described wherein the mixture of fuels is varied to provide a progressively less dense mixture while providing thrust.
U.S. Pat. No. 5,101,622 discloses a rocket engine capable of operating in two propulsive modes for near earth and low earth orbit operations. It describes a first mode in which the external atmosphere is the source of oxidizer for the fuel. At a high Mach number the engine changes to a second mode which is that of a conventional high performance rocket engine using liquid oxygen carried on the vehicle to oxidize a liquid hydrogen fuel. The engine is said to use common hardware including a liquid hydrogen pump and a combustor nozzle assembly. The mechanism required to match the working fuel and oxidizer flow in both propulsive modes is not disclosed. The engine further includes several turbocompressors to compress air to a delivery pressure of several hundred bars, a series of heat exchangers, and turbopumps, all of which make the engine complicated and expensive to produce and operate and its mass prohibitively high.
A liquid air cycle engine, or LACE, is another example of the propulsion were one of the propellant components, in this case oxidizer, can be changed during flight. Liquid oxygen used on the main acceleration mode can be completely or partially replaced by liquefied ambient air during the first propulsive mode beginning at the initial launch or take off through acceleration from sea level atmospheric conditions to moderate speed and altitude. Such an engine is shown and described in H. Hirakoso, “A Concept of LACE for Space Plane to Earth Orbit,” Int. J. Hydrogen Energy, Vol. 15, No. 7, pp. 495-505, at p. 499, 1990.
There is a clear intention in the LACE concepts to maximize the air condensation ratio in an effort to achieve maximum specific impulse or Isp. No real attention is paid, however, to complications to the engine resulting from the necessary additional pumps, plumbing, valves, etc. As a result, the LACE shows inadequate performance gain and/or prohibitively complicated and heavy design. None of the known LACE descriptions suggest the mechanism to match gas flow through the nozzle throat in both the combined and the rocket modes.
Fuel storage systems for rocket engines are shown in U.S. Pat. No. 5,804,760, and U.S. Pat. No. 5,705,771.
The present invention can be based upon existing rocket engines using an expander cycle (RL10 of Pratt& Whitney), gas generator cycle (J2 of Boeing-Rocketdyne), tap-off cycle (J2S, RS2000 of Boeing-Rocketdyne), or a staged combustion topping cycle (SSME of Boeing-Rocketdyne) known in the art. Various examples of rocket engines can be found in D. Huzel and D. Huang, “Modern Engineering for Design of Liquid-Propellant Rocket Engines,” Volume 147 of AIAA Series “Progress in Astronautics and Aeronautics,” pages 35, 36, (1992).
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Crandell Ralph F.
Gartenberg Ehud
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