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Piezo-resonance igniter and ignition method for propellant liquid rocket engine




Title: Piezo-resonance igniter and ignition method for propellant liquid rocket engine.
Abstract: An ignition system for a rocket engine utilizes the pressure energy in a propellant flow. The propellant flow generates an oscillating pressure force in a resonance system which is then transmitted to a piezoelectric system. The electrical pulses are utilized to generate a spark in an igniter system spark gap, resulting in ignition. Since the spark energy production is driven by the resonance of the propellant flow, a fully passive auto-ignition system is provided. Once ignition occurs, the resultant backpressure in the combustion chamber “detunes” the resonance phenomena and spark production stops. Furthermore, should the engine flame out, spark production would automatically resume as the propellant valves remain open thereby providing relight capability. ...


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USPTO Applicaton #: #20090173321
Inventors: Mark D. Horn, Thomas M. Walczuk


The Patent Description & Claims data below is from USPTO Patent Application 20090173321, Piezo-resonance igniter and ignition method for propellant liquid rocket engine.

BACKGROUND

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OF THE INVENTION

The present invention relates to a piezo-resonance igniter system for passive auto ignition of a rocket engine, and more particularly to a method which utilizes the pressure energy in the propellants themselves to excite piezoelectric crystals such that high voltage electrical pulses are created to generate a spark in an igniter system.

Various conventional ignition systems have been used for ignition of a propellant mixture in a combustion chamber of a rocket engine. These ignition systems generally employed a spark induced by an electrical current from a source of electricity and a control for sensing when to supply and discontinue the spark. These conventional systems, although effective, tend to be relatively complex, heavy, and may not provide restart capability.

With the increasing need for safe storable propellant systems such as Gaseous Oxygen (GOx) and Methane combinations, an uncomplicated fully passive auto ignition system is desired to complement the advantages of the safe storable propellants by reducing ignition system complexity, weight, and cost while increasing safety and reliability.

Accordingly, it is desirable to provide an uncomplicated, lightweight passive auto ignition system with restart capability that eliminates separate spark exciter electronics, vehicle electrical power requirements for ignition and ignition control and monitoring systems.

SUMMARY

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OF THE INVENTION

The ignition system according to the present invention generally includes a resonance system in communication with a propellant system, a piezoelectric system, and an electrical conditioning system to power an igniter and ignite a propellant flow. The resonance system is in communication with the piezoelectric system through a gas resonance tube which is sealed with a force transmission diaphragm. A force transfer member increases the surface area in contact with the force transmission diaphragm to react pressure loads from an oscillating flow within the resonance tube. The sizing of the diaphragm allows the resonance pressure pulses to act over a relatively large effective area to increase a net force output for a given resonance gas resonance tube diameter and supply pressure.

The oscillating pressure force from the oscillating flow is transmitted to the piezoelectric crystal stack to generate electrical pulses and power the ignition system. The oscillating force can provide for direct spark ignition in which each pressure pulse results in a spark. Alternatively, the electrical pulses generated may be stored then metered out at various schedules to provide the desired spark repetition rate and spark power per pulse.

Since the spark energy production is driven by the resonance of the propellant flow, a passive auto-ignition system is provided. When the propellant valves are opened, flow through the resonance system is such that resonance occurs and spark energy is generated. Once ignition occurs, the resultant backpressure in the combustion chamber reduces the pressure drop across the resonance system and “detunes” the resonance phenomena, such that spark production stops. Furthermore, should the engine flame out causing combustion chamber pressure to drop again, spark production automatically resumes as long as the propellant valves remain open. Control and operation of the rocket engine is considerably simplified through elimination of the heretofore necessity of an electrical power supply and separate switching commands and monitoring of the ignition system such that the otherwise typical uncertainties in setting spark timing and duration are obviated. Significant advantages are thereby provided for distributed multi-thruster systems, such as an Attitude Control System (ACS), where the characteristics of a conventional ignition system are multiplied by a significant number of thrusters.

The present invention therefore provides an uncomplicated, lightweight passive auto ignition system with restart capability that eliminates separate spark exciter electronics and switching command systems.

BRIEF DESCRIPTION OF THE DRAWINGS

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The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the currently preferred embodiment. The drawings that accompany the detailed description can be briefly described as follows:

FIG. 1 is a general perspective view an exemplary of rocket engine embodiment for use with the present invention;

FIG. 2A is a schematic view of an ignition system of the present invention;

FIG. 2B is an expanded view of the ignition system components illustrated in FIG. 2A;

FIG. 3 is a schematic view of a flight ready ignition system of the present invention with an indirect piezo-resonance module;

FIG. 4 is a schematic view of a flight ready piezo-resonance module ignition system of the present invention utilizing a direct spark torch approach mounted directly within a combustion chamber; and

FIG. 5 is a schematic view of a flight ready piezo-resonance module ignition system of the present invention for use with an incompressible fluid flow.

DETAILED DESCRIPTION

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OF THE PREFERRED EMBODIMENT

FIG. 1 illustrates a general schematic view of a rocket engine 10. The engine 10 generally includes a thrust chamber assembly 12, a fuel system 14, an oxidizer system 16 and an ignition system 18. The fuel system 14 and the oxidizer system 16 preferably provide a gaseous propellant system of the rocket engine 10, however, other propellant systems such as liquid will also be usable with the present invention.

A combustion chamber wall 20 about a thrust axis A defines the nozzle assembly 12. The combustion chamber wall 20 defines a thrust chamber 22, a combustion chamber 24 upstream of the thrust chamber 22, and a combustion chamber throat 26 therebetween. The thrust chamber assembly 12 includes an injector 12A with an injector face 28 which contains a multitude of fuel/oxidizer injector elements 30 (shown somewhat schematically) which receive fuel which passes first through the fuel cooled combustion chamber wall 20 fed via fuel supply line 14a of the fuel system 14 and an oxidizer such as Gaseous Oxygen (GOx) through an oxidizer supply line 16a of the oxidizer system 16.

The ignition system 18 generally includes a resonance system 36 in communication with one of the propellants such as the oxidizer system 16, a piezoelectric system 38, and an electrical conditioning system 40 to power an igniter 42 mounted within the injector 12A to ignite the fuel/oxidizer propellant flow from the fuel/oxidizer injector elements 30. The oxidizer is fed to the igniter via a dedicated line 16b in this embodiment, and the fuel is also fed to the igniter torch via a dedicated line 14b. It should be understood that various propellant flow paths may be usable with the present invention so long as at least one propellant flow is in communication with the resonance system 36. Ignition of the fuel/oxidizer propellant flow from the fuel/oxidizer injector elements 30 with the igniter 42 is conventional and need not be described in further detail herein. It should also be understood that while the current focus of this invention is a rocket ignition, other applications for power generation and ignition of other combustion based devices will also be usable with the present invention.

Referring to FIG. 2A, one ignition system 18 includes a housing 32 which defines a resonance cavity 44 having an inlet 34 incorporating a supersonic inlet nozzle 46 to receive a flow of propellant such as the oxidizer from the oxidizer supply line 16c of the oxidizer system 16. An outlet 16a from the resonance system 36 includes an outlet nozzle 50 to maintain pressure in the cavity 44 at a predetermined level. Although the illustrated embodiment of the oxidizer is a gaseous propellant (compressible flow) resonance configuration, it should be understood that resonant pressure pulses from incompressible liquid flow as well as from other propellant sources will likewise be usable with the present invention.

The resonance system 36 is in communication with the piezoelectric system 38 through a gas resonance tube 52. It should be understood that in FIG. 2A the piezoelectric system 38 is illustrated in a schematic form in what may be considered a ground based configuration which may include adjustment features that may or may not be required. That is, other even less complicated piezoelectric systems are achievable as illustrated in the following embodiments.

The gas resonance tube 52 is located through an opening 54 in the resonance cavity 44 opposite the supersonic inlet nozzle 46. The oxidizer entering through the supersonic inlet nozzle 46 as underexpanded flow is directed at the gas resonance tube 52 causing an oscillating detached shock 56 to form upstream of the entrance 56N to the gas resonance tube 52. Reflected shocks within the gas resonance tube 52 couple and reinforce the detached shock 56 and interact with the flow within the gas resonance tube 52 such that the successive cycles of shocks cause the formation of a series of unstable zones of elevated pressure within the gas resonance tube 52. Physical criteria for the interaction may be defined by: “d” the diameter of the supersonic inlet nozzle 46N; “G” the distance between the inlet nozzle 46N throat and the entrance 56N of the gas resonance tube 52; “Dtube” the internal diameter of resonance tube 52 and “DMC” which is the throat diameter of the outlet nozzle 50. A constant diameter resonance tube 52 is depicted; however, it is understood that stepped, conical or other shaped resonance tubes may alternatively be utilized with the present invention.

The gas resonance tube 52 is sealed at an end opposite the entrance 56N with a force transmission diaphragm 58 (also illustrated in FIG. 2B). A force transfer member 60 includes a force transfer rod 62 and a force transfer platen 64 in contact with the force transmission diaphragm 58. The force transfer platen 64 is of a larger diameter than the force transfer rod 62 so as to increase the surface area in contact with the force transmission diaphragm 58 and react pressure loads from the oscillating pressures in the resonance tube 52. The sizing of the force transmission diaphragm 58 allows the resonance pressure pulses to act over a relatively large effective area, increasing the net force output for a given gas tube 52 diameter (Dtube) and supply pressure. Flow relief passages 52a (FIG. 2B) may be incorporated into the mating faces of the resonance tube 52 and the force transmission diaphragm 58 to increase working fluid transfer across the face of the force transmission diaphragm 58 during the relatively short resonant pressure pulses in the resonance tube 52.

The force transfer rod 62 is received within a guide sleeve 65. The guide sleeve 65 contains a piezoelectric crystal stack 66 mounted in contact with the force transfer rod 62. The oscillating pressure force in the gas resonance tube 52 is transmitted to the piezoelectric crystal stack 66 through the force transfer member 60 to generate electrical pulses. The wire harness 67 is connected directly to the igniter 42, eliminating the electrical conditioning system 40. The oscillating force drives the direct spark ignition, in which each pressure pulse results in a spark, offering a persistent source of ignition.

Alternatively or in addition thereto, the electrical pulses are communicated to the igniter 42 through a wire harness 67 and the electrical conditioning system 40. An energy storage system 68A (illustrated schematically) such as an electrical capacitor or battery and a voltage multiplier system 68B (illustrated schematically) within the electrical condition system 40 conditions the spark to a desired spark output energy and frequency independent of the crystal output. This permits the system to be sized to suit any application. In other words, the electrical condition system 40 may include various electrical subsystems such as storage capacitors or voltage amplifiers to specifically tailor the ignition system to provide various outputs.

Since the spark energy production is driven by the resonance of the propellant flow, a fully passive auto-ignition system is provided. When the propellant valves are open, flow through the resonance system 36 is such that resonance occurs and spark energy is created. Once ignition occurs, the resultant backpressure within the combustion chamber 24 (FIG. 1) “detunes” the resonance phenomena and spark production stops. Furthermore, should the engine flame out, spark production automatically resumes as the propellant valves remain open. Control and operation of the rocket engine is considerably simplified by the elimination of separate power supply and switching command systems in the igniter system such that the heretofore typical uncertainties in the spark duration control are obviated. This provides significant advantages for distributed multi-thruster systems, such as an attitude control system (ACS).

Referring to FIG. 2B, the force transmission diaphragm 58 is preferably sandwiched between an end segment 70 of the gas resonance tube 52 and a diaphragm support ring 72 which may be welded together through a weld W or other attachment. The force transmission diaphragm 58 preferably includes a relief feature 74 located between the diaphragm support ring 72 and the force transfer platen 64. The relief feature 74 is preferably a circular flexed portion of the force transmission diaphragm 58 within which the force transfer platen 64 is received. The relief feature 74 minimizes tensile load losses on the force transmission diaphragm 58 thereby enhancing flexibility to maximize transfer of the oscillating pressure force to the force transfer platen 64 and thence to the piezoelectric crystal stack 66 through the force transfer rod 62.

Applicant has demonstrated relatively short ignition delay times of approximately 18 mseconds utilizing a gaseous propellant (compressible flow) resonance configuration. However, multiple approaches exist to achieve the resonant pressure pulses from incompressible liquid flow as well such that the present invention is adaptable to any propellants.

Referring to FIG. 3, another ignition system 18B is illustrated. The resonance system 36A includes a more compact flight-ready piezoelectric system 38A integrated with the resonance system 36A. Such a system is readily mounted anywhere within the communicating conduits of a working fluid system such as embodied by the oxidizer system or fuel system (FIG. 1).




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stats Patent Info
Application #
US 20090173321 A1
Publish Date
07/09/2009
Document #
File Date
12/31/1969
USPTO Class
Other USPTO Classes
International Class
/
Drawings
0


Liquid Rocket Rocket Engine

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United Technologies Corporation


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Internal-combustion Engines   High Tension Ignition System   Piezoelectric Voltage Generator  

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20090709|20090173321|piezo-resonance igniter and ignition propellant liquid rocket engine|An ignition system for a rocket engine utilizes the pressure energy in a propellant flow. The propellant flow generates an oscillating pressure force in a resonance system which is then transmitted to a piezoelectric system. The electrical pulses are utilized to generate a spark in an igniter system spark gap, |United-Technologies-Corporation
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