US2637161A - Process of ignition for rockets and the like - Google Patents

Process of ignition for rockets and the like Download PDF

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Publication number
US2637161A
US2637161A US126658A US12665849A US2637161A US 2637161 A US2637161 A US 2637161A US 126658 A US126658 A US 126658A US 12665849 A US12665849 A US 12665849A US 2637161 A US2637161 A US 2637161A
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fuel
rocket
combustion
oxidizer
zone
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US126658A
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Johann G Tschinkel
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US Department of Army
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US Department of Army
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/95—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by starting or ignition means or arrangements

Definitions

  • This invention relates to a process for burning fuel and more particularly to a process for energizing rocket motors supplied with liquid fuel and liquid oxidizer.
  • An object of the invention is to provide a process for initiating and continuing combustion in a combustion zone that does not require any auxiliary igniting device, that assures positive and safe functioning, that is simple to operate and that requires a minimum of supplies.
  • a process for burning fuel which comprises bringing together in a combustion zone an oxidizer and a fuel spontaneously combustible therewith to initiate combustion and thereafter supplying to the combustion zone an oxidizer and another fuel to continue combustion.
  • the process finds particular application in the energization of rocket motors, in which use the fuel tank of the rocket is charged with a layer of an initiating fuel and a layer of bulk fuel of a different specific gravity and substantially immiscible therewith, the fuels are passed in a continuous stream to the combustion chamber of the rocket,
  • the initiating fuel entering the chamber first followed by the bulk fuel, and simultaneously there is introduced into the combustion chamber and mixed with the fuel therein a continuous stream of an oxidizer that first reacts spontaneously with the initiating fuel to start combustion and thereafter reacts with the bulk fuel to continue combustion.
  • the single figure is a somewhat diagrammatic view of a portion of a rocket, a part of the shell being broken away to show the interior.
  • the rocket shell I houses, inter alia, a rocket motor or combustion chamber 2 closed at the front 3 and having a reaction nozzle l opening to the rear.
  • Liquid oxidizer 5 contained in oxidizer tank 8 is fed through pipe I under control of motor operated valve 8 to the combustion chamber 2.
  • liquid fuel contained in fuel tank or reservoir 9 is fed to the combustion chambe 2 through pipe Ill under control of motor operated valve l l.
  • Compressed gas, from a source carried by the rocket may be applied to the surfaces of the liquids in the tanks 6 and 9 to force the liquids from the tanks into the combustion chamber.
  • Such gas may enter the tanks through pipes 12 and it having control motor operable valves l4 and Hi respectively.
  • Valve operating motors l8 and [9 may be remotely controlled by well known mechanisms.
  • the drawing represents a layer of initiating fuel and H represents a supernatant layer of bulk fuel.
  • the initiating fuel is substantially immiscible with the bulk fuel and has a density greater than the density of the bulk fuel, as a result of which the fuels separate into two layers as shown.
  • the initiating fuel it and the oxidizer 5 are so selected that they ignite spontaneously when they are mixed in the combustion chamber even at ordinary temperatures. Thus, no preheating or compression is required to ignite the mixture of initiating fuel and oxidizer and no auxiliary igniting device is needed.
  • Sumcient initiating fuel is charged into the fuel tank 9 to start combustion and to heat the combustion chamber to normal operating temperature. After the initiating fuel has been expended, it is immediately followed by the bulk fuel ll which continues to burn at operating temperature. Enough liquid oxidizer 5 is stored in the tank 5 to effect combustion of all of the fuel, both initiating and bulk, contained in tank 9. Ordinarily, the amount of initiating fuel will be less than 5% of the amount of the bulk fuel.
  • the initiating fuel and the bulk fuel may be stored and transported in the same tank or in separate tanks. Where one tank is used for both fuels, separate discharge lines enable each fuel to be separately drained from the tank.
  • the rocket usually is filled in a vertical position, or at least is kept in a vertical position for some time after filling and before firing, there is sufiicient time for substantially complete separation of the fuel into two layers to take place,
  • the oxidizer may be employed as a two-phase system. in which the first phase. to reach the combustion chamber is hypergolic or spontaneously combustible with the first increment of fuel to reach the combustion chamber and the second phase or bulk of the oxidizer is non-hypergolic with the first phase of the fuel.
  • a process for burning fuel in a rocket motor having one combustion zone comprising the steps of bringing together in the combustion zone an oxidizer selected from the group consisting of nitric acid, nitrogen tetroxide and hydrogen peroxide and a fuel spontaneously combustible therewith selected from the groupconsistingbf aniline, furfuryl alcohol, pyrocatechol and hy-'- drazine hydrate to initiate combustion, and thereafter supplying to the combustion zone said oxidizer and another fuel selected from the group consisting of gasoline and benzene to continue combustion.
  • an oxidizer selected from the group consisting of nitric acid, nitrogen tetroxide and hydrogen peroxide and a fuel spontaneously combustible therewith selected from the groupconsistingbf aniline, furfuryl alcohol, pyrocatechol and hy-'- drazine hydrate to initiate combustion, and thereafter supplying to the combustion zone said oxidizer and another fuel selected from the group consisting of gasoline and benzene to continue combustion.
  • a process for energizing a rocket motor having a combustion zone and a fuel'reservoir zone which comprises the steps of charging into the fuel reservoir zone of the rocket a layer of initiating fuel and a supernatant layer of bulk fuel, withdrawing the initiating and bulk fuel from adjacent thebottom of the reservoir zone in a continuous stream, introducing said continuous stream of fuel into the combustion zone ofthe rocket, and simultaneously introducing into the combustion zone of the rocket an oxidizer in an amount and ata rate sufficient to initiate and support combustion of the fuel.
  • a process for energizing a rocket motor having a combustion zone and a'fuel reservoir zone which comprises the steps of charging into the fuel reservoir zone of the rocket a layer of initiating fuel and a supernatant layer of bulk fuel, of lesser density and substantially immiscible therewith, withdrawing the initiating and bulk fuel from adjacent the bottom of the'reser voir zone in a continuous stream, introducing said continuous stream of fuel into the combustion zone of the rocket, and simultaneously introducing into the combustion zone of the rocket an oxidizer in an amount and at a rate sufficient to initiate and support combustion of the fuel.
  • a process for energizing a rocket motor including a combustion zone and a fuel reservoir zone which comprises the steps of charging into the fuel reservoir zone of the rocket a first layer of initiating fuel, superimposing thereon a second, supernatant layer of bulk fuel, withdrawing successively the first and. second layers of fuel from adjacent the bottom of the reservoir zone in a continuous stream, introducing said continuous stream of fuel into the combustion zone of the rocket, and simultaneously introducing into the combustion zone of the rocket an oxidizer in an amount and at a rate suflicient to initiate and support combustion of the fuel.
  • a process for energizing a rocket motor including an oxidizer reservoir zone, a fuel reservoir zone and a combustion zone which comprises the steps of charging into the oxidizer reservoir zone of the rocket a liquid oxidizer, charging into the fuel reservoir zone of the rocket a first layer of initiating fuel spontaneously combustible with said oxidizer at ordinary temperatures, superimposing thereon a second, supernatant layer of bulk fuel spontaneously combustible with said oxidizer only at elevated temperatures, withdrawing successively the first .and second layers of fuel from adjacent the bottom of the reservoir zone in a continuous stream,..introducing .said continuousstream of fuel into the combustion zone of the rocket,.and simultaneously introducing into the combustion zone of the rocket an oxidizer in an amount and at a rat suflicient .to initiate and sustain combustion of the fuel.
  • a processfor energizing-a rocket ,motor including an oxidizer reservoir zone, a fuel reservoir zone and a combustion zone which comprises the steps of charging. into the oxidizer reservoir zone of the rocket a liquid oxidizer, charging into the fuelreservoir zone.
  • a firstlayer of initiating fuelspontaneously combustible with said oxidizer at ordinary temperatures superimposing thereon a second, supernatant layer of bulk fuel immiscible with said initiating fuel and spontaneously combustible with said oxidizer only at elevated temperatures, withdrawing successively the first and second layers .of fuel from adjacent the-bottom of the reservoir zone in a continuous stream, introducing said continuous stream of fuel into the combustion zone of the rocket, and simultaneously introducing into the combustion zone of the rocket an oxidizer in an amount and at a rate suificient to initiate and sustain combustion of the fuel.

Description

May 5, 1953 J. G. TSCHINKEL PROCESS OF IGNITION FOR ROCKETS AND THE LIKE Filed' NOV. 10 1949 L\ qum Oxmnze Q L a u F K L Q Fuel.
OMBUSTION CHAMBER grwwwlom duhnn'n ELTsch'inkel Patented May 5, 1953 orrc PROCESS OF IGNITION FOR ROCKETS AND THE LIKE Application November 10, 1949, Serial No. 126,658
(Granted under Title 35, U. S. Code (1952),
see. 266) 6 Claims.
The invention described herein may be manufactured and used by or for the Government for governmental purposes without th payment of any royalty thereon.
This invention relates to a process for burning fuel and more particularly to a process for energizing rocket motors supplied with liquid fuel and liquid oxidizer.
An object of the invention is to provide a process for initiating and continuing combustion in a combustion zone that does not require any auxiliary igniting device, that assures positive and safe functioning, that is simple to operate and that requires a minimum of supplies.
These and other objects and advantages of the invention, as will appear more fully hereinafter, are realized in a process for burning fuel which comprises bringing together in a combustion zone an oxidizer and a fuel spontaneously combustible therewith to initiate combustion and thereafter supplying to the combustion zone an oxidizer and another fuel to continue combustion. The process finds particular application in the energization of rocket motors, in which use the fuel tank of the rocket is charged with a layer of an initiating fuel and a layer of bulk fuel of a different specific gravity and substantially immiscible therewith, the fuels are passed in a continuous stream to the combustion chamber of the rocket,
the initiating fuel entering the chamber first followed by the bulk fuel, and simultaneously there is introduced into the combustion chamber and mixed with the fuel therein a continuous stream of an oxidizer that first reacts spontaneously with the initiating fuel to start combustion and thereafter reacts with the bulk fuel to continue combustion.
In the drawing, the single figure is a somewhat diagrammatic view of a portion of a rocket, a part of the shell being broken away to show the interior.
Referring to the drawing, the rocket shell I houses, inter alia, a rocket motor or combustion chamber 2 closed at the front 3 and having a reaction nozzle l opening to the rear. Liquid oxidizer 5 contained in oxidizer tank 8 is fed through pipe I under control of motor operated valve 8 to the combustion chamber 2. Similarly, liquid fuel contained in fuel tank or reservoir 9 is fed to the combustion chambe 2 through pipe Ill under control of motor operated valve l l. Compressed gas, from a source carried by the rocket (not shown) may be applied to the surfaces of the liquids in the tanks 6 and 9 to force the liquids from the tanks into the combustion chamber.
Such gas may enter the tanks through pipes 12 and it having control motor operable valves l4 and Hi respectively. Valve operating motors l8 and [9 may be remotely controlled by well known mechanisms.
As seen in the drawing, it represents a layer of initiating fuel and H represents a supernatant layer of bulk fuel. The initiating fuel is substantially immiscible with the bulk fuel and has a density greater than the density of the bulk fuel, as a result of which the fuels separate into two layers as shown. The initiating fuel it and the oxidizer 5 are so selected that they ignite spontaneously when they are mixed in the combustion chamber even at ordinary temperatures. Thus, no preheating or compression is required to ignite the mixture of initiating fuel and oxidizer and no auxiliary igniting device is needed.
Sumcient initiating fuel is charged into the fuel tank 9 to start combustion and to heat the combustion chamber to normal operating temperature. After the initiating fuel has been expended, it is immediately followed by the bulk fuel ll which continues to burn at operating temperature. Enough liquid oxidizer 5 is stored in the tank 5 to effect combustion of all of the fuel, both initiating and bulk, contained in tank 9. Ordinarily, the amount of initiating fuel will be less than 5% of the amount of the bulk fuel.
Effective fuel and oxidizer combinations for use in the process of the invention are listed in the following table:
These are merely exemplary and many other combinations satisfying the conditions outlined herein are available.
The initiating fuel and the bulk fuel may be stored and transported in the same tank or in separate tanks. Where one tank is used for both fuels, separate discharge lines enable each fuel to be separately drained from the tank.
Since the rocket usually is filled in a vertical position, or at least is kept in a vertical position for some time after filling and before firing, there is sufiicient time for substantially complete separation of the fuel into two layers to take place,
'- 3 even though the fuels may have become mixed to some extent durin filling of the tank.
It is also contemplated that the oxidizer may be employed as a two-phase system. in which the first phase. to reach the combustion chamber is hypergolic or spontaneously combustible with the first increment of fuel to reach the combustion chamber and the second phase or bulk of the oxidizer is non-hypergolic with the first phase of the fuel.
I claim: H V
1. A process for burning fuel in a rocket motor having one combustion zone comprising the steps of bringing together in the combustion zone an oxidizer selected from the group consisting of nitric acid, nitrogen tetroxide and hydrogen peroxide and a fuel spontaneously combustible therewith selected from the groupconsistingbf aniline, furfuryl alcohol, pyrocatechol and hy-'- drazine hydrate to initiate combustion, and thereafter supplying to the combustion zone said oxidizer and another fuel selected from the group consisting of gasoline and benzene to continue combustion.
2. A process for energizing a rocket motor having a combustion zone and a fuel'reservoir zone which comprises the steps of charging into the fuel reservoir zone of the rocket a layer of initiating fuel and a supernatant layer of bulk fuel, withdrawing the initiating and bulk fuel from adjacent thebottom of the reservoir zone in a continuous stream, introducing said continuous stream of fuel into the combustion zone ofthe rocket, and simultaneously introducing into the combustion zone of the rocket an oxidizer in an amount and ata rate sufficient to initiate and support combustion of the fuel.
3. A process for energizing a rocket motor having a combustion zone and a'fuel reservoir zone which comprises the steps of charging into the fuel reservoir zone of the rocket a layer of initiating fuel and a supernatant layer of bulk fuel, of lesser density and substantially immiscible therewith, withdrawing the initiating and bulk fuel from adjacent the bottom of the'reser voir zone in a continuous stream, introducing said continuous stream of fuel into the combustion zone of the rocket, and simultaneously introducing into the combustion zone of the rocket an oxidizer in an amount and at a rate sufficient to initiate and support combustion of the fuel.
4. A process for energizing a rocket motor including a combustion zone and a fuel reservoir zone which comprises the steps of charging into the fuel reservoir zone of the rocket a first layer of initiating fuel, superimposing thereon a second, supernatant layer of bulk fuel, withdrawing successively the first and. second layers of fuel from adjacent the bottom of the reservoir zone in a continuous stream, introducing said continuous stream of fuel into the combustion zone of the rocket, and simultaneously introducing into the combustion zone of the rocket an oxidizer in an amount and at a rate suflicient to initiate and support combustion of the fuel.
5. A process for energizing a rocket motor including an oxidizer reservoir zone, a fuel reservoir zone and a combustion zone which comprises the steps of charging into the oxidizer reservoir zone of the rocket a liquid oxidizer, charging into the fuel reservoir zone of the rocket a first layer of initiating fuel spontaneously combustible with said oxidizer at ordinary temperatures, superimposing thereon a second, supernatant layer of bulk fuel spontaneously combustible with said oxidizer only at elevated temperatures, withdrawing successively the first .and second layers of fuel from adjacent the bottom of the reservoir zone in a continuous stream,..introducing .said continuousstream of fuel into the combustion zone of the rocket,.and simultaneously introducing into the combustion zone of the rocket an oxidizer in an amount and at a rat suflicient .to initiate and sustain combustion of the fuel. l
6. A processfor energizing-a rocket ,motor including an oxidizer reservoir zone, a fuel reservoir zone and a combustion zone which comprises the steps of charging. into the oxidizer reservoir zone of the rocket a liquid oxidizer, charging into the fuelreservoir zone. of the rocket a firstlayer of initiating fuelspontaneously combustible with said oxidizer at ordinary temperatures, superimposing thereon a second, supernatant layer of bulk fuel immiscible with said initiating fuel and spontaneously combustible with said oxidizer only at elevated temperatures, withdrawing successively the first and second layers .of fuel from adjacent the-bottom of the reservoir zone in a continuous stream, introducing said continuous stream of fuel into the combustion zone of the rocket, and simultaneously introducing into the combustion zone of the rocket an oxidizer in an amount and at a rate suificient to initiate and sustain combustion of the fuel.
JOHANN G. TSCHINKEL.
References Cited in the file of this patent UNITED STATES PATENTS Great Britain July 10, 1947
US126658A 1949-11-10 1949-11-10 Process of ignition for rockets and the like Expired - Lifetime US2637161A (en)

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Cited By (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2699385A (en) * 1951-04-19 1955-01-11 Gulf Research Development Co Hydrocarbon fuel oil compositions
US2753683A (en) * 1952-06-24 1956-07-10 Standard Oil Co Liquid propellant
US2873577A (en) * 1955-05-09 1959-02-17 Gen Electric Combustion system for jet engine starters
US2880582A (en) * 1956-04-19 1959-04-07 Clement J Turansky Starting assembly for a power plant
US2939278A (en) * 1955-02-28 1960-06-07 Phillips Petroleum Co Means and method for starting rocket motors
US2974484A (en) * 1952-01-23 1961-03-14 Robert A Cooley Ignition system for rocket motors
US2993334A (en) * 1954-08-02 1961-07-25 Phillips Petroleum Co Ignition delay reducing agents for hypergolic rocket fuels
US3024598A (en) * 1958-02-17 1962-03-13 Berliner Maschb Ag Vorm L Schw Hydrojet engine for marine and submarine propulsion
US3058301A (en) * 1958-12-24 1962-10-16 Phillips Petroleum Co Reaction motor fuels
US3075463A (en) * 1959-09-04 1963-01-29 Dow Chemical Co Well fracturing
US3088274A (en) * 1959-10-19 1963-05-07 Thompson Ramo Wooldridge Inc Dual thrust rocket engine
US3128601A (en) * 1960-09-15 1964-04-14 United Aircraft Corp Pre-burner rocket control system
US3212254A (en) * 1957-08-14 1965-10-19 Phillips Petroleum Co Two component amine nitrate monopropellants and method of propulsion
US6393830B1 (en) * 1999-03-26 2002-05-28 Alliant Techsystems Inc. Hybrid rocket propulsion system including array of hybrid or fluid attitude-control rocket engines

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US830144A (en) * 1905-02-21 1906-09-04 Hiram A Frantz Explosive-engine.
US1183038A (en) * 1915-08-16 1916-05-16 Busch Sulzer Bros Diesel Engine Co Operation of engines on tar.
US2408111A (en) * 1943-08-30 1946-09-24 Robert C Truax Two-stage rocket system
GB590177A (en) * 1944-07-17 1947-07-10 Hydran Products Ltd Improvements in or relating to projectiles of the rocket type
US2523008A (en) * 1946-07-26 1950-09-19 Daniel And Florence Guggenheim Turbine starting and control apparatus
US2573471A (en) * 1943-05-08 1951-10-30 Aerojet Engineering Corp Reaction motor operable by liquid propellants and method of operating it

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US830144A (en) * 1905-02-21 1906-09-04 Hiram A Frantz Explosive-engine.
US1183038A (en) * 1915-08-16 1916-05-16 Busch Sulzer Bros Diesel Engine Co Operation of engines on tar.
US2573471A (en) * 1943-05-08 1951-10-30 Aerojet Engineering Corp Reaction motor operable by liquid propellants and method of operating it
US2408111A (en) * 1943-08-30 1946-09-24 Robert C Truax Two-stage rocket system
GB590177A (en) * 1944-07-17 1947-07-10 Hydran Products Ltd Improvements in or relating to projectiles of the rocket type
US2523008A (en) * 1946-07-26 1950-09-19 Daniel And Florence Guggenheim Turbine starting and control apparatus

Cited By (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2699385A (en) * 1951-04-19 1955-01-11 Gulf Research Development Co Hydrocarbon fuel oil compositions
US2974484A (en) * 1952-01-23 1961-03-14 Robert A Cooley Ignition system for rocket motors
US2753683A (en) * 1952-06-24 1956-07-10 Standard Oil Co Liquid propellant
US2993334A (en) * 1954-08-02 1961-07-25 Phillips Petroleum Co Ignition delay reducing agents for hypergolic rocket fuels
US2939278A (en) * 1955-02-28 1960-06-07 Phillips Petroleum Co Means and method for starting rocket motors
US2873577A (en) * 1955-05-09 1959-02-17 Gen Electric Combustion system for jet engine starters
US2880582A (en) * 1956-04-19 1959-04-07 Clement J Turansky Starting assembly for a power plant
US3212254A (en) * 1957-08-14 1965-10-19 Phillips Petroleum Co Two component amine nitrate monopropellants and method of propulsion
US3024598A (en) * 1958-02-17 1962-03-13 Berliner Maschb Ag Vorm L Schw Hydrojet engine for marine and submarine propulsion
US3058301A (en) * 1958-12-24 1962-10-16 Phillips Petroleum Co Reaction motor fuels
US3075463A (en) * 1959-09-04 1963-01-29 Dow Chemical Co Well fracturing
US3088274A (en) * 1959-10-19 1963-05-07 Thompson Ramo Wooldridge Inc Dual thrust rocket engine
US3128601A (en) * 1960-09-15 1964-04-14 United Aircraft Corp Pre-burner rocket control system
US6393830B1 (en) * 1999-03-26 2002-05-28 Alliant Techsystems Inc. Hybrid rocket propulsion system including array of hybrid or fluid attitude-control rocket engines

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