US3540679A

US3540679A - Unified rocket control

Info

Publication number: US3540679A
Application number: US738723A
Authority: US
Inventors: Edward E Mccullough; Donald E Hume
Original assignee: DONALD E HUME
Current assignee: DONALD E HUME
Priority date: 1968-06-13
Filing date: 1968-06-13
Publication date: 1970-11-17
Anticipated expiration: 1987-11-17

Description

United States Patent Inventors Edward E. McCullough;

Donald E. Hume, Brigham City, Utah 84302 Appl. No. 738,723 Filed June 13, 1968 Patented Nov. 17, 1970 Assignee the United States of America as represented by the Secretary of the Air Force UNIFIED ROCKET CONTROL 4 Claims, 2 Drawing Figs.

US. Cl 244/322, 102/494, 60/204 Int. Cl F41g 7/00, F42b 15/02 Field of Search 244/322,

[56] References Cited UNITED STATES PATENTS 2,726,510 12/1955 Goddard.

FOREIGN PATENTS 941,108 6/1948 France.

Primary Examiner-Samuel Feinberg Attorneys-Harry A. Herbert, Jr. and Arsen Tashjian ABSTRACT: A rocket motor for steering a ballistic missile during the powered stage of its flight. The motor, which is of the liquid propellant rocket type, is affixed to the forward end of the ballistic missile and includes a plurality of combustion chambers having angularly oriented! plug nozzles. When it becomes necessary to change the course of the missile, fuel and oxidizer are introduced into one or more of the combustion chambers, to cause the resultant thrust to accomplish the desired change of course.

Patented Nov. 17, 1970 3,540,619

INVENTORJ. mwmw 4'. Mn c'z/uauay,

. 1 'UNIFIED ROCKET CONTROL BACKGROUND OF THE INVENTION This invention relates to the art of steering of ballistic type missiles and, more particularly, to a rocket motor of the liquid propellant type having aplurality of combustion chambers for steeringsuch a'missile during the powered portion of its flight.

The term ballistic missile," as used herein, is intended to mean any missile, including a multiple-stage one, which is launched, follows a selected course, is powered to the desired velocity and desired altitude, and which, during the nonpowered phase, follows a free-flight path consisting essentially of a ballistic trajectory, irrespective of whether the terminal portion of the flight is guided or not.

Steering of a ballistic type missile, during the powered portion of its flight, has been attempted in many ways. All of them are intended primarily to prevent undesirable roll, pitch, and yaw, and, thereby, to provide stabilization of the missile from launch of the entire missile, i.e., stages and payload or warhead, until entry of the payload or warhead into the ballistic trajectory,

Some of the methods and apparatuses include, but are not All of the above-mcntioned methods and apparatuses, as

well as others known but not mentioned. herein, have two things in common. Firstly, they are 'used, or are intended-for use, at or near the aft end of the missile. This results in an undesirably shorter, relatively speaking, lever arm to the center of mass ofthe missile as a whole. Secondly, all add weight and complexity to the missile. The added weight results in a geometric increase of the thrust required to propel the missile. The added complexity increases the probability ofthe unreliability of the missile as a whole. In addition; some of the steering methods and apparatuses may create unique peripheral problems. For example: Where movable nozzles are used, premature erosion may be introduced, and added to, existing problems. 7

All of these'problems'are multiplied and magnified by the fact that, as thestages of the missile increase in number, each stage is provided with an independent steering means, while the sole guiding means remains at the payload or warhead end ofthe missile.

This invention avoids all of these disadvantages, and obviates these problems, by providing a simple, highly reliable, unified steering apparatus for the entire missile, and positioning said steering apparatus at the payload or warhead, notaft. end ofthe missile.

In additionthis invention further advances the state-oftheart by providing aerodynamic protection for the nose cone portion ofthe payload or warhead, by permitting aerodynamic heating ofthepropellant used in the invention, by allowing the use of propellant in the invention for forward acceleration when any of the propellant is not needed for steering, and by increasing the length of the lever arm to the center of mass of the missile.

SUMMARY OF THE INVENTION I This invention, a rocket motor of the liquid propellant type,

A still further object of the invention is to provide a lightweight apparatus for steering ballistic missiles, thereby resulting in a significant saving in weight over other state-ofthe-art steering apparatuses. Other additional, and equally important, objects of this invention are the mechanical advantageof a longer lever arm to the center of mass ofthe ballistic missile, aerodynamic protection for the nose cone proper of the missile, aerodynamic heating of the propellant in the invention, and the lack of consumption of thrust from the main operating engine of the ballistic missile. 1

These, and still other, objects of the invention will become readily apparent after a consideration of the description of the invention and ofthe drawings.

DESCRIPTION OF THE DRAWINGS FIG. I is a side elevation view of'a'multiplestage ballistic missile in a ready-for-launch attitude, with a schematic representation of the invention in position and ready for use, at the forward end ofthe missile. FIG. 2 is a side elevation view, schematic in form, partially in vertical cross section and fragmented in part, of the principal components of the invention, with the said invention in position on the nose cone or payload ofthe missile.

DESCRIP-TIONOF THE PREFERRED EMBODIMENT 7 With reference to FIGS. 1 and 2, wherein the same components are similarly numbered, the invention, a liquid propellant rocket motor 20, is attached (by means notshown) to the forward end, i'.e., the nose cone ll, of ballistic type missile 10.

In FIG. 1, missile 10 is shown, asa matter ofpreference, as a multiple-stage one. Also shown is the center ofmass 12 of missile l0.

With reference to FIG. 2, liquid propellant rocket motor 20 has two separatecompartments, one 21 is for the fuel and the other 22 is for the oxidizer. The wall 23 which separates these compartmentsZl and 22 contains a valve 24 for metering the desired quantity of the fuel and of the oxidizer into any one or more of the four combustion chambers 25 for ultimate expulsion through throats 26, which are equipped with plug nozzles which are permanently rearwardly oriented at approximately 45 to the axis ofmissile 10.

MODE OF OPERATION OF THE EMBODIMENT When it becomes necessary to change the course of missile 10, fuel, such as hydrazine, from compartment 21, and an oxidizer, such as hydrogen peroxide, from compartment 22, are

are already firing, changes the course of missile 10 to the desired one.

When missile 10 has reached the end of the powered phase ofits flight, and is about to enter the free-flight ballistic trajec' tory portion ofits flight, rocket motor 20 is separated, by suitable remote means, from the nose cone 11 and, simultaneously, all four combustion chambers 25 are tired to assist in the separation.

It is to be noted that, although rocket motor 20 described in the preferred embodiment is of the liquid propellant type, a rocket motor ofthe hybrid type, may be used.

Further, although rocket motor 20 is described, in the preferred embodiment, as having four combustion chambers 25, there may bemore, or less, of such chambers 25 in a particular embodiment to achieve the samepurpose.

While there has been shown and described the fundamental features of our invention, as'applied. to a preferred embodi ment, it is understood that various substitutions and omissions, such as described above, may be made by those skilled in the art without departing from the spirit ofthe invention.

We claim:

3 a 4 l. The' 'method-of steering a'missile of the ballistic type comprising the steps of:

a. mountinga detachable multiple chamber rocket motor on the forward portion of the ballistic missile;

b. firing at least one chamber of'said rnultip'le chamber rocket motor, to accomplish 'on course positioning of the ballistic missile during the powered portion of its flight;

2. The'methodt as set forth in claim 1, wherein the multiple chamber rocket motor is of'the liquid propellant type.

v 3; The method, as set forth in claim 1, wherein the multiple.

chamber rocket motor is of the hybrid propellant type.

4. A rocket motor for steering a missile of the ballistic type,

comprising;

a. a casing c onfigurated at the aft end to abut upon the nose cone of the missile and dctachably mounted upon the acompa'rtment. withintsaid casing, to hold a liquid rocket fuel; m

. a compartment, within said casing, to hold a liquid rocket oxidizer; l

..a plurality'of combustion chambers, within said casing,

and equally spaced from each other;

. means, leading from said liquid fuel compartment to said comb'usti'onchambers, for metering said liquid fuel to said combustion chambers;

. means leading' from said liquid oxidizer to'said comw bustion chambers,'for metering said liquid mtidizer to said combustion chambers} nozzles, permanently rearwardly oriented at approximately 45 to the. axis of the missile leading from said combustion chambers to the exterior otsaid casing; and means'for detaching, 'byvremote control said detachably mounted casing and its contents.