US3178885A - Hybrid rocket engine - Google Patents

Hybrid rocket engine Download PDF

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US3178885A
US3178885A US116416A US11641661A US3178885A US 3178885 A US3178885 A US 3178885A US 116416 A US116416 A US 116416A US 11641661 A US11641661 A US 11641661A US 3178885 A US3178885 A US 3178885A
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nozzle
liquid
oxidizer
solid
component
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US116416A
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Robert H Loughran
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Lockheed Corp
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Lockheed Aircraft Corp
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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
    • F02K9/72—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid and solid propellants, i.e. hybrid rocket-engine plants

Definitions

  • This invention relates generally to hybrid rocket motors, which term is used to describe rocket motors that use separate fuel and oxidizer components, one of which is liquid and the other solid. More particularly, the invention is concerned with a rocket motor of this class which vaporizes and heats a liquid oxidizer to a high temperature before contact with the solid fuel.
  • One of the main objects of the invention is to provide a solid propellant ignition heater and gas generator that pressurizes a liquid oxidizer and vaporizes Tes'ime as the heater exhausts through an underexpanded flow nozzle upon the reaction surface of the solid fuel.
  • Another object of the invention' is' to provide an ignition heater that pressurizes a liquid oxidizer and vaporizes the same at a flow nozzle for high temperature diffusion over the solid fuel react iye surface.
  • Still another object of the invention is to provide an igniter that pressurizes, vaporizes and heats a liquid oxidizer to high temperature before contact with the solid fuel.
  • a further object of the invention is to provide an ignition heater for hybrid rocket engines, which vaporizes and heats a liquid oxidizer before the latter contacts a solid fuel, and thus permits the use of various combinations of solid fuel and liquid oxidizer, that are not normally reactive when the oxidizeris in the liquid phase.
  • a still further object of the invention is to provide an ignition heater, the high temperature combustion products of which change a liquid oxidizer to a gaseous condition and also break down the fuel composition to more reactive.
  • Yet another object of the invention is to provide an ignition heater and gas generator which vaporizes and preheats a liquid component, so that many combinations of liquid and solid fuel and oxidizer may be used.
  • Still another object of the invention is to provide a relatively simple and low cost type of injector that insures hot gas phase reaction and high burning rate and at the same time has capability of Omi oif operation ⁇ n rmal sw n-
  • FIGURE 1 is a schematic view of a hybrid rocket engine, showing a preferred form of the invention.
  • a solid fuel rocket motor casing of conventional design has the usual exhaust nozzle 11 at the rear end thereof. Bonded to the inner surface 13 of the casing 10 is a body 12 of solid fuel which may, if desired, have some oxidizer incorporated therein.
  • the body 12 of solid fuel has a central, longitudinally extending bore 14 which forms the burning surface.
  • the bore 14 may be of circular cross-section, or it may take any other conmake it figuration.
  • the elements heretofore recited are entirely conventional.
  • the solid propellant 17 differs from the solid fuel 12 of the motor casing 10 in one important respect, and that is that it comprises a fuel and oxidizer mixture in proportions for stoichiometric combustion, whereas the fuel 12 is either completely without oxidizer, or with a substantial deficiency of oxidizer, whereby combustion of the fuel 12 cannot be supported without additional oxidizer.
  • the solid propellant 17 is provided with a central longitudinally extending bore 18 from one end of the casing 16 to the other, the rear end of said bore opening into an under-expanded nozzle 19.
  • the nozzle 1-9 is attached to the front end of the motor casing 10 around the marginal edges of an aperture 20, and the nozzle 19 thus discharges directly into the bore 14.
  • An igniter 21 is mounted on the front end of the casing 16, and serves to ignite the solid propellant 17 when fired.
  • the combustion gases from the auxiliary gas generator 15 are exhausted through the nozzle 19 into the rocket engine and upon the reaction surface 14 of the fuel 12.
  • a pressure tube 22 is connected at 23 to the auxiliary gas generator 15, and at 24 to a container 25 filled with highly reactive liquid oxidizer 26, such as chlorine trifluoride (ClF for example.
  • a conventional piston 30 with standard sealing rings 31 is disposed in the pressure end 32 of the casing 15.
  • One end of a supply tube 33 is connected at 34 to the container 25, and its other end communicates with a plurality of ports, or discharge openings 35, in the throat of the nozzle 19.
  • a valve 36 is provided in the supply tube 33 for controlling the flow of liquid oxidizer, and thereby controlling the rate of combustion of the fuel 12, which has the effect of controlling the propulsive thrust of the rocket motor 10.
  • Maximum rate of combustion of the fuel 12 for maximum thrust is obtained when the quantity of oxidizer 26 discharged into the bore 14 is exactly equal to the requirement for stoichiometric combustion.
  • the thrust of the rocket motor 10 may be programmed in any desired manner by merely operating the valve 36 to restrict the flow of oxidizer. If desired, combustion of the fuel 12 may be stopped by closing the valve 36, and even restarted by opening the valve again, provided the auxiliary gas generator 15 is still operating.
  • Operation of the invention is effected by firing the igniter 21, which ignites the small solid propellant 17 in the auxiliary gas generator 15.
  • the combustion gases of the auxiliary gas generator are discharged through the under-expanded flow nozzle 19 into the bore 14.
  • the nozzle 19 effectively becomes a heater nozzle due to the hot gases passing therethrough.
  • the combustion gases of the generator 15 pressurize the liquid oxidizer through the tube 22, acting against the piston 30.
  • the liquid oxidizer 26 controlled by the valve 36 is forced through the throat holes 35 in the heater nozzle 19.
  • the heater nozzle gas flow vaporizes and heats the liquid oxid'uer to high temperature before contact with the solid fuel 12.
  • the vaporized liquid oxidizer is diffused completely within the rocket engine and over the reactive surface.
  • the invention is characterized by insensitivity to scaling, high burning rate, and high mass ratio. There is also an insensitivity of specific impulse to ratio of fuel and oxidizer.
  • the rocket engine is rendered flexible and enables the use of combinations of fuel and oxidizer that are not normally reactive when the oxidizer is in a liquid state. The hot gaseous condition of the oxidizer tends to break down the fuel composition to make it more reactive.
  • Various combinations of conventional solid fuels, oxidized and slightly oxidized, or solid fuels alone may be used.
  • the present invention contemplates the use of combinations of solid oxidizers such as ammonium nitrate, for example, and liquid fuel, such as gasoline or kerosene.
  • solid oxidizers such as ammonium nitrate, for example
  • liquid fuel such as gasoline or kerosene.
  • the solid oxidizer would be loaded in the motor case 10, and the liquid fuel would be carried in the container 25.
  • a hybrid rocket engine comprising a motor case having a rearwardly directed exhaust nozzle, a body of solid fuel contained within said motor case, said body of solid fuel defining a combustion port opening at one end into said exhaust nozzle, an auxiliary gas generator having a nozzle discharging into said combustion port at the other end thereof, a container of liquid oxidizer, a conduit opening at one end into said container and at the other end into said last-named nozzle, and a second conduit connecting said auxiliary gas generator with said container, whereby gas pressure developed in said generator is applied to said container, and forces said liquid oxidizer into said last-named nozzle where it is vaporized before being discharged into said combustion port.
  • SAMUEL LEVINE Primary Examiner.
  • ABRAM BLUM Examiner.

Description

fiEAHCH HUGE/'1 A ril 20, 1965 R. H. LOUGHRAN HYBRID ROCKET ENGINE Filed June 12, 1961 IN V EN TOR. Poss/2r H. LOUGHRAA/ United States Patent 3,178,885 HYBRID ROCKET ENGINE Robert H. Loughran, Redlands, CaHf., assignor, by mesne assignments, to Lockheed Aircraft Corporation, Burbank, Calif., a corporation of California Filed June 12, 1961, Ser. No. 116,416 3 Claims. (Cl. 6035.6)
This invention relates generally to hybrid rocket motors, which term is used to describe rocket motors that use separate fuel and oxidizer components, one of which is liquid and the other solid. More particularly, the invention is concerned with a rocket motor of this class which vaporizes and heats a liquid oxidizer to a high temperature before contact with the solid fuel.
One of the main objects of the invention is to provide a solid propellant ignition heater and gas generator that pressurizes a liquid oxidizer and vaporizes Tes'ime as the heater exhausts through an underexpanded flow nozzle upon the reaction surface of the solid fuel.
Another object of the invention'is' to provide an ignition heater that pressurizes a liquid oxidizer and vaporizes the same at a flow nozzle for high temperature diffusion over the solid fuel react iye surface.
Still another object of the invention is to provide an igniter that pressurizes, vaporizes and heats a liquid oxidizer to high temperature before contact with the solid fuel.
A further object of the invention is to provide an ignition heater for hybrid rocket engines, which vaporizes and heats a liquid oxidizer before the latter contacts a solid fuel, and thus permits the use of various combinations of solid fuel and liquid oxidizer, that are not normally reactive when the oxidizeris in the liquid phase.
A still further object of the invention is to provide an ignition heater, the high temperature combustion products of which change a liquid oxidizer to a gaseous condition and also break down the fuel composition to more reactive.
Yet another object of the invention is to provide an ignition heater and gas generator which vaporizes and preheats a liquid component, so that many combinations of liquid and solid fuel and oxidizer may be used.
Still another object of the invention is to provide a relatively simple and low cost type of injector that insures hot gas phase reaction and high burning rate and at the same time has capability of Omi oif operation \n rmal sw n- The foregoing and other objects and advantages of this invention will be clear to those skilled in the art upon consideration of the following detailed specification of one illustrative embodiment thereof, reference being had to the accompanying drawing, in which:
FIGURE 1 is a schematic view of a hybrid rocket engine, showing a preferred form of the invention.
Referring to the drawing, which is more or less schematic, it will be noted that the various elements are not shown in complete structural detail. This has not been deemed necessary because it is believed that a profusion of drawings would merely complicate the disclosure of fundamental elements and the combination thereof. The prior art recognizes certain essential elements and obviates the necessity of engineering detail.
A solid fuel rocket motor casing of conventional design has the usual exhaust nozzle 11 at the rear end thereof. Bonded to the inner surface 13 of the casing 10 is a body 12 of solid fuel which may, if desired, have some oxidizer incorporated therein. The body 12 of solid fuel has a central, longitudinally extending bore 14 which forms the burning surface. The bore 14 may be of circular cross-section, or it may take any other conmake it figuration. The elements heretofore recited are entirely conventional.
Mounted on the front end of the motor casing 10 is an auxiliary gas generator 15 comprising a casing 16 loaded with a grain of solid propellant 17. The solid propellant 17 differs from the solid fuel 12 of the motor casing 10 in one important respect, and that is that it comprises a fuel and oxidizer mixture in proportions for stoichiometric combustion, whereas the fuel 12 is either completely without oxidizer, or with a substantial deficiency of oxidizer, whereby combustion of the fuel 12 cannot be supported without additional oxidizer. The solid propellant 17 is provided with a central longitudinally extending bore 18 from one end of the casing 16 to the other, the rear end of said bore opening into an under-expanded nozzle 19. The nozzle 1-9 is attached to the front end of the motor casing 10 around the marginal edges of an aperture 20, and the nozzle 19 thus discharges directly into the bore 14. An igniter 21 is mounted on the front end of the casing 16, and serves to ignite the solid propellant 17 when fired. The combustion gases from the auxiliary gas generator 15 are exhausted through the nozzle 19 into the rocket engine and upon the reaction surface 14 of the fuel 12.
A pressure tube 22 is connected at 23 to the auxiliary gas generator 15, and at 24 to a container 25 filled with highly reactive liquid oxidizer 26, such as chlorine trifluoride (ClF for example. A conventional piston 30 with standard sealing rings 31 is disposed in the pressure end 32 of the casing 15. One end of a supply tube 33 is connected at 34 to the container 25, and its other end communicates with a plurality of ports, or discharge openings 35, in the throat of the nozzle 19. A valve 36 is provided in the supply tube 33 for controlling the flow of liquid oxidizer, and thereby controlling the rate of combustion of the fuel 12, which has the effect of controlling the propulsive thrust of the rocket motor 10. Maximum rate of combustion of the fuel 12 for maximum thrust is obtained when the quantity of oxidizer 26 discharged into the bore 14 is exactly equal to the requirement for stoichiometric combustion. The thrust of the rocket motor 10 may be programmed in any desired manner by merely operating the valve 36 to restrict the flow of oxidizer. If desired, combustion of the fuel 12 may be stopped by closing the valve 36, and even restarted by opening the valve again, provided the auxiliary gas generator 15 is still operating.
Operation of the invention is effected by firing the igniter 21, which ignites the small solid propellant 17 in the auxiliary gas generator 15. The combustion gases of the auxiliary gas generator are discharged through the under-expanded flow nozzle 19 into the bore 14. Thus the nozzle 19 effectively becomes a heater nozzle due to the hot gases passing therethrough. At the same time, the combustion gases of the generator 15 pressurize the liquid oxidizer through the tube 22, acting against the piston 30. The liquid oxidizer 26 controlled by the valve 36, is forced through the throat holes 35 in the heater nozzle 19. The heater nozzle gas flow vaporizes and heats the liquid oxid'uer to high temperature before contact with the solid fuel 12. The vaporized liquid oxidizer is diffused completely within the rocket engine and over the reactive surface. The invention is characterized by insensitivity to scaling, high burning rate, and high mass ratio. There is also an insensitivity of specific impulse to ratio of fuel and oxidizer. The rocket engine is rendered flexible and enables the use of combinations of fuel and oxidizer that are not normally reactive when the oxidizer is in a liquid state. The hot gaseous condition of the oxidizer tends to break down the fuel composition to make it more reactive. Various combinations of conventional solid fuels, oxidized and slightly oxidized, or solid fuels alone may be used.
In addition to the solid fuel and liquid oxidizer combination described herein, the present invention contemplates the use of combinations of solid oxidizers such as ammonium nitrate, for example, and liquid fuel, such as gasoline or kerosene. In this instance, the solid oxidizer would be loaded in the motor case 10, and the liquid fuel would be carried in the container 25.
While I have shown and described in considerable detail what I believe to be the preferred form of my invention, it will be understood by those skilled in the art that various changes may be made in the shape and arrangement of the several parts without departing from the broad scope of the invention as defined in the following claims.
I claim:
1. A hybrid rocket engine utilizing afuel component and an oxidizer component, one of said components being solid and the other being liquid, said engine comprising a motor case loaded with said solid component and having a rearwardly directed nozzle through which combusliquid component is contained, an auxiliary ga ge r 1e .=1. .9l f
mounted on the front end of the motor case and liavlng a nozzle discharging into said motor case, means for pressurizing said container with gas generated by said auxiliary gas generator, means for conveying said liquid component under pressure from said container to said lastnamed nozzle and means including said last-named nozzle for mixing said liquid component with the hot combustion gases discharged by said auxiliary gas generator, whereby the said liquid component is vaporized as it is discharged into said motor case for reaction with said solid component.
2. A hybrid rocket engine comprising a motor case having a rearwardly directed exhaust nozzle, a body of solid fuel contained within said motor case, said body of solid fuel defining a combustion port opening at one end into said exhaust nozzle, an auxiliary gas generator having a nozzle discharging into said combustion port at the other end thereof, a container of liquid oxidizer, a conduit opening at one end into said container and at the other end into said last-named nozzle, and a second conduit connecting said auxiliary gas generator with said container, whereby gas pressure developed in said generator is applied to said container, and forces said liquid oxidizer into said last-named nozzle where it is vaporized before being discharged into said combustion port.
3. A hybrid rocket engine as defined in claim 2, wherein said auxiliary gas generator comprises a second case loaded with solid propellant, and means for igniting said solid propellant.
References Cited by the Examiner UNITED STATES PATENTS 2,753,801 7/56 Cumming 35.6 2,791,883 5/57 Moore et al. 6039.47 2,878,643 3/59 Fox 6035.6 2,940,256 6/60 Conyers et al 60--39.48 2,972,225 2/61 Cumming et a1. 603948 2,984,973 5/61 Stegelman 6039.48 2,987,875 6/61 Fox 6035.6 2,996,880 8/61 Greiner 6035.6 2,998,703 9/61 Badders 6035.6 3,017,748 1/62 Burnside 6035.6 3,065,597 '9 11/62 Adamson et al 6035.6 2,065,598 11/62 Schultz 6035.6 3,068,641 12/62 Fox 6035.6 3,115,007 12/63 Fox 6035.6
SAMUEL LEVINE, Primary Examiner. ABRAM BLUM, Examiner.

Claims (1)

1. A HYBRID ROCKET ENGINE UTILIZING A FUEL COMPONENT AND AN OXIDIZER COMPONENT, ONE OF SAID COMPONENTS BEING SOLID AND THE OTHER BEING LIQUID, SAID ENGINE COMPRISING A MOTOR CASE LOADED WITH SAID SOLID COMPONENT AND HAVING A REARWARDLY DIRECTED NOZZLE THROUGH WHICH COMBUSTION PRODUCTS ARE DISCHARGED, A CONTAINER IN WHICH SAID LIQUID COMPONENT IS CONTAINED, AN AUXILIARY GAS GENERATOR MOUNTED ON THE FRONT END OF THE MOTOR CASE AND HAVING A NOZZLE DISCHARGING INTO SAID MOTOR CASE, MEANS FOR PRESSURIZING SAID CONTAINER WITH GAS GENERATED BY SAID AUXILIARY GAS GENERATOR, MEANS FOR CONVEYING SAID LIQUID COMPONENT UNDER PRESSURE FROM SAID CONTAINER TO SAID LASTNAMED NOZZLE AND MEANS INCLUDING SAID LAST-NAMED NOZZLE FOR MIXING SAID LIQUID COMPONENT WITH THE HOT COMBUSTION GASES DISCHARGED BY SAID AUXILIARY GAS GENERATOR, WHEREBY THE SAID LIQUID COMPONENT IS VAPORIZED AS IT IS DISCHARGING INTO SAID MOTOR CASE FOR REACTION WITH SAID SOLID COMPONENT.
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Cited By (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3334489A (en) * 1964-06-04 1967-08-08 Propulsion Par Reaction Sa Soc Rocket motor
US3421323A (en) * 1966-11-14 1969-01-14 Donald Perry Bennett Jr Fluid fuel and non-fluid oxidizer energy generation method
US3439612A (en) * 1966-11-14 1969-04-22 United Aircraft Corp Hybrid flare
US3518828A (en) * 1968-09-27 1970-07-07 Us Air Force Hybrid rocket motor ignition system
US3777490A (en) * 1972-03-10 1973-12-11 Nasa Supersonic-combustion rocket
US3782112A (en) * 1972-02-24 1974-01-01 Us Navy Hybrid generator
US3908358A (en) * 1973-01-31 1975-09-30 Thiokol Corp Variable flow gas generating method and system
US5099645A (en) * 1990-06-21 1992-03-31 General Dynamics Corporation, Space Systems Division Liquid-solid propulsion system and method
US5274998A (en) * 1992-07-06 1994-01-04 Wyle Laboratories Rocket pollution reduction system
US5339625A (en) * 1992-12-04 1994-08-23 American Rocket Company Hybrid rocket motor solid fuel grain
US5794435A (en) * 1996-02-07 1998-08-18 Lockhhed Martin Corporation Stable-combustion oxidizer vaporizer for hybrid rockets
US6073437A (en) * 1994-10-13 2000-06-13 Lockheed Martin Corporation Stable-combustion oxidizer for hybrid rockets
US20030136107A1 (en) * 2002-01-22 2003-07-24 Hy Pat Corporation Hybrid rocket motor having a precombustion chamber
US6629673B2 (en) * 2001-11-28 2003-10-07 United Technologies Corporation Adaptable solid-hybrid rocket for crew escape and orbital injection propulsion
US20060145018A1 (en) * 2003-03-28 2006-07-06 Rutan Elbert L Unitized hybrid rocket system
WO2013048271A1 (en) 2011-09-29 2013-04-04 Omnidea Lda. Propulsion system
WO2018217264A1 (en) * 2017-02-28 2018-11-29 Alpha Space Test And Research Alliance Llc Multi-stage solid rocket motor

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* Cited by examiner, † Cited by third party
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US2065598A (en) * 1936-04-10 1936-12-29 Aaron E Mccoy Tractor
US2753801A (en) * 1952-02-28 1956-07-10 James M Cumming Combination liquid and solid propellent rocket
US2791883A (en) * 1951-10-25 1957-05-14 Gen Electric Propellant system
US2878643A (en) * 1955-05-09 1959-03-24 Phillips Petroleum Co Combustion stabilization control system responsive to oxidant concentration
US2940256A (en) * 1954-03-26 1960-06-14 North American Aviation Inc Ullage compensation for pressurizing systems
US2972225A (en) * 1950-12-04 1961-02-21 James M Cumming Motor mechanism for missiles
US2984973A (en) * 1958-12-08 1961-05-23 Phillips Petroleum Co Liquid-solid bipropellant rocket
US2987875A (en) * 1955-05-26 1961-06-13 Phillips Petroleum Co Ramjet power plants for missiles
US2996880A (en) * 1958-10-14 1961-08-22 Texaco Experiment Inc Reaction propulsion system and rocket
US2998703A (en) * 1953-09-11 1961-09-05 William C Badders Reso-jet igniter
US3017748A (en) * 1959-01-02 1962-01-23 Phillips Petroleum Co Combination liquid and solid propellant spin-stabilized rocket motor
US3065597A (en) * 1959-09-28 1962-11-27 Gen Electric Reignitable solid rocket motor
US3068641A (en) * 1955-04-18 1962-12-18 Homer M Fox Hybrid method of rocket propulsion
US3115007A (en) * 1958-09-22 1963-12-24 Phillips Petroleum Co Self-actuating hybrid rocket motor

Patent Citations (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2065598A (en) * 1936-04-10 1936-12-29 Aaron E Mccoy Tractor
US2972225A (en) * 1950-12-04 1961-02-21 James M Cumming Motor mechanism for missiles
US2791883A (en) * 1951-10-25 1957-05-14 Gen Electric Propellant system
US2753801A (en) * 1952-02-28 1956-07-10 James M Cumming Combination liquid and solid propellent rocket
US2998703A (en) * 1953-09-11 1961-09-05 William C Badders Reso-jet igniter
US2940256A (en) * 1954-03-26 1960-06-14 North American Aviation Inc Ullage compensation for pressurizing systems
US3068641A (en) * 1955-04-18 1962-12-18 Homer M Fox Hybrid method of rocket propulsion
US2878643A (en) * 1955-05-09 1959-03-24 Phillips Petroleum Co Combustion stabilization control system responsive to oxidant concentration
US2987875A (en) * 1955-05-26 1961-06-13 Phillips Petroleum Co Ramjet power plants for missiles
US3115007A (en) * 1958-09-22 1963-12-24 Phillips Petroleum Co Self-actuating hybrid rocket motor
US2996880A (en) * 1958-10-14 1961-08-22 Texaco Experiment Inc Reaction propulsion system and rocket
US2984973A (en) * 1958-12-08 1961-05-23 Phillips Petroleum Co Liquid-solid bipropellant rocket
US3017748A (en) * 1959-01-02 1962-01-23 Phillips Petroleum Co Combination liquid and solid propellant spin-stabilized rocket motor
US3065597A (en) * 1959-09-28 1962-11-27 Gen Electric Reignitable solid rocket motor

Cited By (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3334489A (en) * 1964-06-04 1967-08-08 Propulsion Par Reaction Sa Soc Rocket motor
US3421323A (en) * 1966-11-14 1969-01-14 Donald Perry Bennett Jr Fluid fuel and non-fluid oxidizer energy generation method
US3439612A (en) * 1966-11-14 1969-04-22 United Aircraft Corp Hybrid flare
US3518828A (en) * 1968-09-27 1970-07-07 Us Air Force Hybrid rocket motor ignition system
US3782112A (en) * 1972-02-24 1974-01-01 Us Navy Hybrid generator
US3777490A (en) * 1972-03-10 1973-12-11 Nasa Supersonic-combustion rocket
US3908358A (en) * 1973-01-31 1975-09-30 Thiokol Corp Variable flow gas generating method and system
US5099645A (en) * 1990-06-21 1992-03-31 General Dynamics Corporation, Space Systems Division Liquid-solid propulsion system and method
US5274998A (en) * 1992-07-06 1994-01-04 Wyle Laboratories Rocket pollution reduction system
US5339625A (en) * 1992-12-04 1994-08-23 American Rocket Company Hybrid rocket motor solid fuel grain
US6073437A (en) * 1994-10-13 2000-06-13 Lockheed Martin Corporation Stable-combustion oxidizer for hybrid rockets
US5794435A (en) * 1996-02-07 1998-08-18 Lockhhed Martin Corporation Stable-combustion oxidizer vaporizer for hybrid rockets
US6629673B2 (en) * 2001-11-28 2003-10-07 United Technologies Corporation Adaptable solid-hybrid rocket for crew escape and orbital injection propulsion
US20030136107A1 (en) * 2002-01-22 2003-07-24 Hy Pat Corporation Hybrid rocket motor having a precombustion chamber
US6679049B2 (en) * 2002-01-22 2004-01-20 Hy Pat Corporation Hybrid rocket motor having a precombustion chamber
US20040055277A1 (en) * 2002-01-22 2004-03-25 Hy Pat Corporation Hybrid rocket motor having a precombustion chamber
US6820412B2 (en) 2002-01-22 2004-11-23 Hy Pat Corporation Hybrid rocket motor having a precombustion chamber
US20060145018A1 (en) * 2003-03-28 2006-07-06 Rutan Elbert L Unitized hybrid rocket system
US7540145B2 (en) * 2003-03-28 2009-06-02 Mojave Aerospace Ventures, Llc Unitized hybrid rocket system
WO2013048271A1 (en) 2011-09-29 2013-04-04 Omnidea Lda. Propulsion system
WO2018217264A1 (en) * 2017-02-28 2018-11-29 Alpha Space Test And Research Alliance Llc Multi-stage solid rocket motor

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