|Publication number||US3149460 A|
|Publication date||Sep 22, 1964|
|Filing date||Sep 28, 1960|
|Priority date||Sep 28, 1960|
|Publication number||US 3149460 A, US 3149460A, US-A-3149460, US3149460 A, US3149460A|
|Inventors||La Rocca Aldo V|
|Original Assignee||Gen Electric|
|Export Citation||BiBTeX, EndNote, RefMan|
|Patent Citations (13), Referenced by (22), Classifications (9)|
|External Links: USPTO, USPTO Assignment, Espacenet|
Sept 22, 1954 A. v. LA RoccA REACTION PRoPuLsIoN SYSTEM Filed Sept. 28. 1960 United States Patent O 3,149,469 REACTION PROPULSION SYSTEM Aldo V. La Rocca, Cincinna Ghio, assigner to General Eiectric Company, a corporation of New York Fiied Sept. 23, 1969, Ser. No. 58,928 1 Claim. (Cl. 60-35.6)
The present invention relates to a reaction propulsion system and, more particularly to a compounded reaction propulsion system which makes use of three processes of heat addition to a working iluid.
These processes can be used separately or in combination to obtain a range of thrust and specilic impulses when a working uid is expanded through a nozzle for the creation of thrust. In such a manner a device better capable of matching the propulsive requirements for several applications is obtained, resulting in an improved overall performance.
In any reaction propulsion system, it is desirable to have high equivalent specific impulse. This is expressed as:
Ff Wes The equivalent specific impulse (I) can thus be defined as the product of the thrust (F) times the propulsion time (tf) which product is the total impulse, all divided by the weight of the propulsion system (Wps). In the more general case where the thrust is not constant the equivalent specic impulse is expressed by the integral of the thrust over the propulsion time divided by the propulsion system weight and expressed as:
tf =in Fai To maximize the equivalent specic impulse i.e. to optimize the propulsiva system for a given application, particular values of the speciiic impulse are required. Specic impulse is defined as the exhaust jet velocity (Vex) divided by the standard acceleration of gravity (go) thus:
Vex I a,
These values of the specific impulse can be greater than those possible with chemical systems i.e. those systems making use only of the energy stored in the chemical bonds of the propellants or generated in an exothermic reaction such as oxidation or combustion.
Other means of achieving high specific impulses for systems discharging to a vacuum can be obtained by raising the settling chamber temperature to levels above those possible by a chemical reaction process. Another possible means is that of adding energy in a form other than thermal to the working fluid. Of the irst class of devices the electrical me reaction engine makes use of addition of thermal energy at a very hiUh temperature to a working iiuid. The range of settling chamber temperature depends on the dilution of the arc by the amount of the working fluid injected at a given level of electrical power. The usable stagnation temperatures are limited basically by two phenomena, (l) the containment problem or the structural and mechanical integrity of the chamber and of the electrodes which have to withstand very large heat iluxes. And (2) the occurrence of chemical processes such as dissociation and physical processes such as changes of state and ionization. Both of these would impair the eiiiciency of the device, if the associated energies could not be recovered and converted to thrust in the exhaust nozzle.
Examples of the second class of devices are the plasma 3,149,460 Patented Sept. 22., 1964 accelerator and the ion accelerator. They increase the kinetic energy of the jet by electromagnetic or electro static forces. These, on the other hand, face several development problems. Furthermore their intrinsic characteristics seem to indicate advantageous performance only for propulsive requirements resulting in optimizing speciiic impulses several times larger than those of chemical systems. The gap that is left in the specific impulse between the chemical and the second class of devices just described offers a specific impulse range which is required to optimize propulsive systems for several applications of earth orbiting vehicles. It is this gap in which the propulsion system of the present invention is designed to operate.
The main object of the present invention is to provide a reaction propulsion system that will deliver a wide range of thrust variation with highest equivalent specific impulse and with a minimum of Variable geometry and control complexity.
A further object is to provide such a system which does not have the storage, temperature and cooling problems of the normal high specific electrothermal engines.
Another object is to provide such a system wherein thrust augmentation is obtainable without any variable geometry.
Briefly stated, the invention provides a reaction propulsion system with a heating chamber, electrical energy means in the heating chamber for heating the fuel, a source of exothermally decomposable fuel of the type which, when it decomposes produces products of a molecular weight of less than 16, and a converging-diverging nozzle downstream of the heating chamber to discharge the heated products of decomposition to produce thrust. A further modiiication provides a thrust augmenting means by which an additional reactant such as an oxidizing agent is introduced into a combustion chamber to raise the temperature of the decomposed products and provide higher thrust by discharging through a second converging-diverging nozzle downstream of the first nozzle.
While the specification concludes with a claim particularly pointing out and distinctly claiming the subject matter which l regard as my invention, it is believed the invention will be better understood from the following description taken in connection with the accompanying drawing in which:
The drawing is a schematic sketch of a reaction propulsion system of the present invention.
Referring to the drawing, the system consists of a heating chamber 1t?. Suitable storage means 11 supplies an exotherrnally decomposable compound to the chamber 10 through line 12 under control of valve 13. For illustrative purposes such a compound may be diborane, pentaborane, hydrazine, lithium hydrate or nitro-methane.
These compounds have the advantage of easy, stable and compact storage in the liquid form. Thus, the problems associated with the extremely low temperature (cryogenic) storage of some other proposed working iiuid such as hydrogen and helium are avoided. Furthermore, the exothermic reaction of decomposition does away with the need for supplying the heat of vaporization which is required by other possible Working fluids such as water. Such a reaction gives the added advantage that heat is released in the decomposition in an amount that can be up to 15% of the total heat that must be added to the working iluid.
The decomposition process may be spontaneous at the conditions existing in the heating chamber 10 or can be assured by appropriate decomposer devices 14 between the storage means 11 and the throat of the nozzle which may be a glow plug, a catalytic surface or the equivalent.
In the instant invention it is important that the fuel source 11 contain fuel which is exothermally decomposable and which is decomposable into products having a molecular weight of not more than fifteen.
These products represent a good Working fiuid for the propulsive device either by direct expansion to vacu um or withtwo further heat addition processes.
One of these processes provides heat to the decomposed fuel in chamber by a suitable electrical energy discharge means which is provided preferably, although not necessarily, centrally of chamber 1f). This electrical energy discharge means may be conveniently an electric arc plasma generator, which is nothing more than a direct or alternating electrical current discharge of high intensity that heats the propellant, Which is directed by suitable means at the end of line 12 to surround the arc, and adds energy to it. The term electrical energy discharge means is intended to include any such electric arc of this general type whether or not associated with electrical preheating schemes such as the more conventional electrical heaters of the resistance or inductance type. The electric arc may be suitably poW- ered by the usual external means indicated generally at 16 for heating the decomposed products to a temperature far above those obtainable by chemical reaction.
The kproducts of decomposition are heated in this manner to appropriately high temperatures compatible with the above mentioned containment and efficiency criteria. The'heated products are then expanded (and thereby accelerated) through a first converging-diverging nozzle 17 into a combustion chamber 18 which surrounds the first chamber preferably concentrically. The exhaust of 'this first nozzle is further or continuously expanded to the external ambient (vacuum) by means of a second convergingt-diverging nozzle 19 which Vcan be considered as an extension of the first as far as theexpansion process is concerned, although structurally it is attached to the combustion chamber.
The propulsion system thus far described, by using an exothermally dec'ornposable fuel that decomposes into products of decomposition having a molecular weight of not more than fifteen, permits an engine of a high specific impulse without the disadvantages of the high temperature, high Weight features and 'permits easy storability since the monopropellant is dense and easily storable under ordinary conditions.
A second feature of the instant invention is to provide augmented thrust to the propulsion system just described. This is carried out by another heat addition process in which an appropriate reactant` such as an oxidizer in container 20 is fed to chamber 18, by means of a yline 21 and control valve 22. The oxidizer is directed to regeneratively cool the internally placed chamber 10 and nozzle 17 and, by mixing it with the products of decomposition, which may or may not have been further heated by the electrothermal device 15 in chamber 10, further heat is added to the Working fluid and the thrust is augmented. A
For the augmentation process the first nozzle 17 will behave as a torch or fiame holder insuring sure starting, better mixing, and more eii'icient combustion. It can be seen that, in this configuration, the second nozzle 19 is an extension of the 'iirst and preferably concentrically to provide a tandem arrangement. The advantages of this arrangement are several:
'(1) The first 17 Vand second nozzles 19 essentially represent a single nozzle of very good area ratio for the products of the first two heat addition processes which are decomposition and electrothermal heat addition. Furthermore, an aerodynamic or self-adjusting effect automatically ensures the proper attachment of the exhaust jet of the first nozzle 17 to the inner surface -of the nozzle 19 at 23 'for the best fiuidynamic efiiciency during operation of the two heat addition processes, namely, decomposition and electrothermal heat addition.
(2) Small amounts of a reactant (such as an oxidizer) or worlc'ng fiuid can be bled through line 21 or a bypass portion 24 of line 12 under control of valve 25 respectively to regeneratively cool the first chamber 10 and nozzle 17 when needed. With the configuration shown, the cooling fluid is then mixed, burned to increase the temperature in the second chamber and expanded at high efficiency Vtaking advantage of the self-adjusting feature mentioned above.
(3) When the reactant or oxidizer fiowv is increased to give full augmentation, the appropriately sized and located throat 26 of the second nozzle 19 downstream of the exit of the first nozzle controls the flow.v
Thus, the tandem nozzle configuration actually performs as a variable nozzle without requiring movable hardware, and will be self-adjusting for any mode of operation.
While have hereinbefore described a preferred form of my invention, obviously many modications and variations of the present invention are possible in the light of the above teachings. It is therefore to be understood that, Within the scope of the appended claim, the invention may be practiced otherwise than as specifically described.
A reaction propulsion system comprising, in combination:
a heating chamber;
electrical high temperature energy discharge means disposed in the said chamber;
a source of exothermally decomposable fuel of the type which upon decomposition produces products of a molecular Weight less than 16;
means connecting the said chamber and exothermally decomposable fuel source and directing the fuel into said chamber for exothermal decomposition and for heating by said energy means, after decomposition, to temperatures far above those obtainable by chemical reaction;
a converging-diverging nozzle connected to said heating chamber downstream thereof to expand and accelerate the heated products of decomposition;
a combustion chamber surrounding the exit from the said nozzle;
a source of oxidizer;
means connecting the said source of oxidizer with the said combustion chamber to introduce oxidizer therein to oxidize the expanded accelerated decomposition products emerging 'om the said nozzle to further increase their energy and to increase total mass flow;
second nozzle means connected to the said combustion chamber to discharge, expand, and further accelerate the said decomposition products and oxidizer.
References Cited in the file of this patent UNITED STATES PATENTS 2,532,709 Goddard Dec. 5, 1950 2,648,190 Maisner Aug. ll, 1953 2,648,317 Mikulasek Aug. ll, 1953 2,706,887 GroW Apr. 276, 1955 2,751,750 Welch et al June 26, 1956 2,753,801 Cumming July 10, 1956 2,775,866 Randall Ian. 1, 1957 2,850,662 Gilruth et al. Sept. 2, 1958 2,887,844 Coty May 26, 1959 2,952,122 Fox Sept. 13, 1960 2,981,059 Horner et a1. Apr. 25, 1961 3,016,693 Jack et al. Jan. 16, 1962 3,041,824 Berhman July 3, 1962 OTHER REFERENCES Rocket Propulsion Elements, by Sutton, 1949, pages -102.
|Cited Patent||Filing date||Publication date||Applicant||Title|
|US2532709 *||Nov 30, 1946||Dec 5, 1950||Daniel And Florence Guggenheim||Liquid cooled baffles between mixing and combustion chambers|
|US2648190 *||Mar 5, 1948||Aug 11, 1953||Aerojet General Co||Initiation of propellant decomposition|
|US2648317 *||Feb 2, 1948||Aug 11, 1953||Mikulasek Libor||Operation of combustion motors by hydrazine|
|US2706887 *||Jan 23, 1946||Apr 26, 1955||Grow Harlow B||Liquid propellant rocket motor|
|US2751750 *||Oct 19, 1954||Jun 26, 1956||British Thomson Houston Co Ltd||Reaction chambers for the decomposition of monofuels|
|US2753801 *||Feb 28, 1952||Jul 10, 1956||Cumming James M||Combination liquid and solid propellent rocket|
|US2775866 *||Sep 10, 1954||Jan 1, 1957||British Thomson Houston Co Ltd||Starters for prime movers such as gas turbines|
|US2850662 *||Mar 4, 1958||Sep 2, 1958||Gilruth Robert R||Electric arc powered jet|
|US2887844 *||May 17, 1952||May 26, 1959||Coty Fred P||Rocket motor|
|US2952122 *||Apr 29, 1955||Sep 13, 1960||Phillips Petroleum Co||Fuel system for ducted rocket ramjet power plants|
|US2981059 *||Feb 4, 1958||Apr 25, 1961||Thompson Ramo Wooldridge Inc||Dual thrust chamber rocket|
|US3016693 *||Sep 23, 1960||Jan 16, 1962||Jack John R||Electro-thermal rocket|
|US3041824 *||May 1, 1956||Jul 3, 1962||Amalgamated Growth Ind Inc||Propulsion system|
|Citing Patent||Filing date||Publication date||Applicant||Title|
|US3279177 *||Jun 10, 1963||Oct 18, 1966||Giannini Scient Corp||Apparatus and method for propelling vehicles in space|
|US3304719 *||Jul 28, 1964||Feb 21, 1967||Giannini Scient Corp||Apparatus and method for heating and accelerating gas|
|US3350887 *||Mar 6, 1964||Nov 7, 1967||Wasagchemie Ag||Two-stage rocket propulsion system|
|US3451221 *||Jul 26, 1966||Jun 24, 1969||Marquardt Corp||Supersonic combustion nozzle|
|US3533233 *||Sep 13, 1967||Oct 13, 1970||Lockheed Aircraft Corp||Hot gas generator utilizing a mono-propellant fuel|
|US3595022 *||Apr 9, 1969||Jul 27, 1971||Licentia Gmbh||Thermodynamic reaction drive|
|US3651644 *||Jun 25, 1969||Mar 28, 1972||Marshall Ind||Apparatus for initiating decomposition of an exothermic propellant|
|US3695041 *||May 8, 1970||Oct 3, 1972||Rocket Research Corp||Two-stage hydrazine rocket motor|
|US3719046 *||Jul 2, 1970||Mar 6, 1973||Rocket Research Corp||Rocket engine cooling system|
|US3956885 *||Sep 3, 1974||May 18, 1976||Avco Corporation||Electrothermal reactor|
|US4133173 *||Jun 13, 1977||Jan 9, 1979||The United States Of America As Represented By The Secretary Of The Navy||Ducted rockets|
|US4147025 *||Mar 14, 1977||Apr 3, 1979||Vereinigte Flugtechnische Werke-Fokker Gmbh||Formation of auxiliary drive gas for turbine|
|US4730449 *||Dec 19, 1986||Mar 15, 1988||Technion, Inc.||Radiation transfer thrusters for low thrust applications|
|US4825647 *||Jan 24, 1985||May 2, 1989||Technion, Inc.||Performance improvements in thruster assembly|
|US5165227 *||Apr 11, 1991||Nov 24, 1992||Mtu Motoren- Und Turbinen-Union Munchen Gmbh||Propelling nozzle for a hypersonic engine|
|US5282357 *||Apr 19, 1990||Feb 1, 1994||Trw Inc.||High-performance dual-mode integral propulsion system|
|US5417049 *||Aug 21, 1991||May 23, 1995||Trw Inc.||Satellite propulsion and power system|
|US5572865 *||Sep 1, 1994||Nov 12, 1996||Trw Inc.||Satellite propulsion and power system|
|US7389636 *||Jul 6, 2005||Jun 24, 2008||United Technologies Corporation||Booster rocket engine using gaseous hydrocarbon in catalytically enhanced gas generator cycle|
|US20130199155 *||Jan 3, 2013||Aug 8, 2013||Jordin Kare||Rocket Propulsion Systems, and Related Methods|
|US20140182265 *||Jan 3, 2013||Jul 3, 2014||Jordin Kare||Rocket Propulsion Systems, and Related Methods|
|USRE32918 *||Nov 13, 1986||May 9, 1989||Technion, Inc.||Heater/emitter assembly|
|U.S. Classification||60/260, 60/39.462, 60/258, 313/231.31, 60/786|
|International Classification||F02K9/00, F02K9/68|