US 3034434 A
Description (OCR text may contain errors)
May 15, 1962 F. H. swAlM ET AL 3,034,434
THRUsT VECTOR CONTROL SYSTEM.
Filed March 8, 1960 2 Sheets-Sheet 1 W gg/ M ATTORNEYS May 15, 1962 F. H. swAlM ETAL THRUsT VECTOR CONTROL SYSTEM 2 Sheets-Sheet 2 Filed March 8. 1960 xNvEmoRs FRANK H.swA1M FRED A. oYHus @W cwu ATTORNEYS United States Patent Oiilice .y j
3,034,434 THRUST VECTOR CONTROL SYSTEM .Y Frank H. Swaim, Silver Spring, and Fred A. Oyhus, Rockville, Md., assignors to the United States of America as represented by the Secretary of the Navy Filled Mar. 8, 1960, Ser. No. 13,662
1 Claim. (Cl. 102-50) l A further object of the present invention is to provide i an incremental thrust vector control system so constructed as to effectivelysteer guided missiles without causing excessive structural loading of the airframes thereof.
Another object of the present invention is to provide an incremental thrust vector control system employing a plurality of selectively dischargeable rockets to effect steering of guided missiles at high altitudes.
A still further object of this invention is to provide an incremental thrust vector control system that requires a minimum of components and that is therefore economical to construct.
Other objects and many `of the attendant advantages oi this invention Will be readily appreciated as the` same becomes better understood by reference to the following detailed description whenconsidered in connection with the accompanying drawings, wherein:
FIG. l is a plan view of a guided missile employing the incremental thrust vector control system constituting the present invention;
FIG. 2 is a schematic of the incremental thrust vector control system, the steering unit employed being shown partially in section and partially in elevation;
FIG. 3 is an enlarged perspective view of the rocket unit employed;
FIG. 4 is a detail section, on the line 4--4 of FIG. 2, showing the construction of the rocket unit;
FIG. 5 is a perspective View showing the manner of operation of the incremental thrust vector control system; and
FIG. 6 is a diagrammatic view showing the operation of a missile having the present invention incorporated therein.
This invention relates to a thrust vector control system for steering guided missiles at altitudes of the order of 100,000 feet and higher. The density of the atmosphere at such altitudes is such that the steering ability of the normal aerodynamic control surfaces of a guided missile is greatly impaired. It is therefore desirable to employ additional steering means on such missiles, said means preferably operating independently of the atmosphere and serving as either the sole steering means or as an auxiliary steering means to aid the usual aerodynamic control surfaces. It is the purpose of this invention to provide such a high altitude steering system.
The invention contemplates the employment of a plurality of steering rockets disposed in the missile yat a position removed from the center of gravity thereof. The rockets can be ignited individually to incrementally align the thrust vector of the missile propulsion system along a desired course. The steering rockets are arranged in four equally spaced rows about the longitudinal axis of the missile, each such row being in alignment with said axis. In the embodiment of the invention described hereinbelow the steering rockets are disposed in the nose of the missile; but it is to be understood that they could equallywell be disposed in the tail, lor at any position removed from the center of gravity of the missile.
Guided missiles such as are herein employed may lbe launched at movingV targets which ily at altitudes inexcess ofV 100,000 feet. The guidance system of the missile is designed to track such targets and to supply information to themissile steering system whereby the missile is directed toward the target. As previously mentioned, at altitudes above 100,000 feet the density of the atmosphere is such that the steering elect of aerodynamic control surfaces is greatly impaired. When such condition is reached in a missile employing the present invention,
the guidance system of the missile will sense the lack of response to the aerodynamic control surface movements and will thereupon emit a signal which will ignite one or more of the steering rockets, whereby the missile is turned in increments as may be required. When the missile has been turned the desired amount, a second signal from the guidance system will ignite one or morev of the steering rockets disposed on the side opposite said iirst ignited rockets whereby the turning motion of the missile is halted. The corrections made to direct the missile ltoward its target must occur while its propulsion system is operating; otherwise, the missile would not be dithe direction of the thrust vector created by the missile l propulsion system whereby it is brought into alignment with theI target. Once the thrust vector is so aligned, the missile,.willrcontinue on that course until its direction yis again caused to be changed.
AReferring to the drawings, a missile 2 is shown in FIG. l. The nose of said missile includes a steering unit 4, said unit having a plurality of exhaust ports 6 therein, through which steering rockets exert their thrust. .TheV exhaust ports are arranged in rows, each row being in alignment with the longitudinal axis of the missile. According to the invention, the rows of steering rockets are employed in pairs, the rockets of one row of a pair exhausting outwardly in a direction degrees removed from the direction of exhaust of the other row of said pair. Thus, each pair of rows of steering rockets provides for steering in one plane, the rockets of one row of a pair being ignited in sutlicient number to initiate steering action, `and the rockets of the second row of said pair being ignited in suicient number to stop the steering faction and to stabilize the missile on its course. steering in more than one plane is desired, then more than one pair of rows of steering rockets must be employed. While any number could be employed, it has` been rfound that two pairs of rows of ports and their associated steering rockets, the four rows comprising the tWo pairs being equally spaced about the missile to form a cruciiorm, will provide complete steering for a missile. The number of ports in each row is a matter of design .and is dependent upon the degree of maneuverability lrequired and the thrust value of each individual steering rocket, such thrust value being so designed that the incremental direction change caused by ignition of each steering rocket will not unduly strain the airframe of the missile or cause serious over-correction in the course.
An incremental thrust vector control system is shown in FIG. 2, the system including the steering unit 4, shown in cross-section, a switching system 8, and a guidance system 10. The steering unit 4 includes a rocket unit 12 having an outer skinV 14, and bulkheads 16 and 18 for mounting the rocket unit 12 within said skin 14 and for Patented May 15, 196,2
securing the steering unit to a missile, the bulkhead 16 connecting the unit 4 to a nose cone 29 .and the bulkhead 18 being employed to connect the unit 4 to the remainder of the missile structure (not shown). The specic construction and configuration of the bulkheads is a matter of design; one possible construction would be to fabricate the bulkheads from sheet metal. Similarly, the means employed to secure the rocket unit 4, the outer skin 14, the nose cone 2@ and the remainder of the missile structure to the bulkheads is a matter of design; screw fasteners and welding are two possible such means. The'rocket unit 12 consists of a plurality of individual steering rockets 22 -arranged in rows in the manner previously indicated. The construction of the unit 12 will be more fully explained hereinbelow.
The specic constructions of the guidance system 1t) and the switching system 3 are not a part of this invention. Generally, the guidance system consists of the usual missile guidance system, to which has been added a sensing device to determine the point at which the missiles response to the aerodynamic control surfaces fails to attain -a pre-set value. Suitable circuitry is employed to connect the sensing device to the switching system for transmitting a signal from the sensing device to the switching system to activate the same. The switching system is constructed to receive the steering correction signals from the guidance system and, in response thereto, to ignite in proper sequence the number of rockets 22 necessary to give the desired steering corrections. The switching system is constructed so that if one rocket 22 fails to ignite another rocket in the same row will be activated, thereby negating the eects of a malfunctioning rocket 22.
A rocket unit 12 is shown in perspective in FIG. 3. Generally, the unit 12 consists of a cruciform-shaped manifold 24 formed in a stepped manner so as to fit within the generally conical interior of a missile nose (see FlG. 2). Referring to FIG. 4, which is a cross-section, taken generally at 4--4 in FIG. 2, of the rocket unit 12 as installed in a missile nose, the unit 12 is seen to have a plurality of bores 26 into each of which a nozzle 27 is press-fitted, or otherwise secured. An amount of propellant (not shown) necessary to generate the desired thrust is disposed within each bore, the propellant being ignited by suitable means (not shown). Each bore 26, when it contains its propellant, comprises an individual steering rocket 22; each such rocket 22 is connected to the switching system 8 by a wire (not shown) which passes through -a central passage 2S in the manifold 24. As is obvious, individual rockets mounted on a suitable frame could be employed instead of the unitary rocket unit 12. However, the unitary nature of the manifold 24 provides numerous structural and economic advantages. Each rocket 22 is exhausted through a port 6 in the `missile skin. The ports 6 are initially provided with tightly tting covers 30 to smooth aerodynamic iiow lines during the major portion of the flight. The covers 30 are readily removed by the force of the steering rocket exhaust upon rocket ignition.
The operation of the incremental thrust vector control center of gravity disposed at 32, is shown in ight. The
thrust generated by the missile propulsion unit is directed along a vector which passes through the center of gravity 32. When it is desired to change the direction, or course, of the missile to the position shown in dotted lines, steering rockets 22 are ignited in sequence 'to steer the vehicle to the desired course, the individual steering rockets each causing an incremental change in the ldirection of the missile, When the desired course is 'obtained steering rockets in suihcient number to stabilize the missile on said course are ignited.
Referring to FIG. 6, the flight of a missile employing an incremental thrust vector control system is shown in diagrammatic form. A target is indicated at A, the target proceeding toward an intercept point B. The missile 2 is shown at C as it reaches the final period of its powered flight. By such time the missile has attained a very high altitude, and as a result its .aerodynamic control surfaces will not exert control eects suilicient to alter its course to cause it to intercept target A at point B. Accordingly, upon a signal emitted by the guidance `system 1t), individual rockets 22 are ignited, as indicated at 34, to align the missile on -a course which passes through intercept point B. Once the missile is so aligned (indicated at D), more individual rockets 22 are ignited (at 36) to stabilize it on its course, thus placing it on an interception course whereby it will intercept the target A at point B.
Obviously, many modifications 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.'
What is claimed isi:
In combination, a missile, a steering unit disposed in said missile at a pointY removed from the center of gravity thereof, said unitV including an outer skin that forms a portion of the outer skin of said missile and that has two pairs of rows of ports therein, each row of ports being displaced degrees from an adjacent row and being in alignment with the longitudinal axis of said missile, cover means frictionally retained within each of said ports, a rocket unit disposed within said steering unit, said rocket unit including an elongated, unitary, cruciformshaped in cross-section manifold having a plurality of bores therein for the reception of propellant, one of said bores being positioned to confront each of said ports, and an exhaust nozzle secured within each said bore in alignment with the port confronted by said bore.
References Cited in the file of this patent UNITED STATES PATENTS 2,395,435 Thompson et al Feb. 25, 1946 2,415,348 Haigney Feb. 4, 1947 2,726,510 Goddard Dec. 13, 1955 2,956,401 Kane Oct. 18, 1960 2,968,454 Merrill et al. Jan. 17, 1961