US 3067682 A
Description (OCR text may contain errors)
Dec. 11, 1962 F. K. FELDMANN ETAL 3,057,682
GYRO PULL ROCKET 4 Sheets-Sheet 1 Filed Feb. 18. 1960 M NRAv MAME. MARR E DHRT N LLEE H LPv ZZRI. KK WmAl REHH Dec. 1l, 1962 F. K. FELDMANN ETAL GYRO PULL ROCKET 4 Sheets-Sheet 3 Filed Feb. 18. 1960 /VVEWTOS FRITZ K. FELDMNN HEINZ F. GEHLHAAR CHARLES RHERRMNN PHILIP C.PETRE yfg/@adzfgw' l l TT/EYS De@ 11, 1962 F. K. FELDMANN ErAL 3,067,682
GYRo PULL ROCKET Filed Feb. 18. 1960 4 Sheets-Sheet 4 XXXXXXX/VVYYYXXXXXXXXXXXXX' FIG. 7
58 F nr./f" /6' I 60 y L-Lvfi /G n X/VYXXM/V YX XX YV ATTORNEYS 3,067,682 GYRO PULL RCKET Fritz K. Feldmann, Heinz l?.l Gehlhaar, Charles R. Herrmann, and Philip C. Petre, all of Santa Barbara, Calit., assignors, by mesne assignments, to Aerojet-General Corporation, Azusa, Calif., a corporation of Ohio Filed Feb. 18, 1960, Ser. No. 9,649 8 Claims. (Cl. 102-49) This invention relates generally to rocket propelled missiles and more particularly to improved booster rockets in conjunction with either ballistic or self-powered pay loads.
Present day booster rockets are generally of the pusher type in which a payload is launched with the aid of the booster until the booster fuel is exhausted. At this time the booster stage is separated from the pay load. More recently, it has been proposed to employ boosters for controlling the range of a ballistic missile by separating the Ibooster from the pay load when a certain condition is established regardless of whether or not the booster fuel .has been used up. In this way, the pay load itself follows a ballistic trajectory after separation from the booster with a predetermined velocity so that accurate ranging can be achieved.
Present day booster rockets however, are subject to many problems, one of the more serious of which is thrust misalignment. Particularly is this the case in solid propellant rockets in which unsy-mmetrical physical conditions in the rocket motor during burning can result in dispersion of the missile. Therefore, when a solid propellant booster rocket is employed, any thrust misalignment can result in large position errors. In addition, the application of thrust of push boosters is generally behind the center of gravity of the pay load so -that directional stability is often times diiicult to maintain. Thus While some speed control may be realized by the use of push boosters for ballistic type missiles, accurate direction control is not always achieved.
Directional errors as a result of aerodynamic disturbances or other external disturbing forces 0n the misvsile and booster is alsoa serious problem. Heretofore relatively sophisticated solutions such as jet vanes, reaction controls, specialized launchers, and aerodynamic .control surfaces have been proposed but their degree of success is usually proportional to their complexity .and
thus contribute to inherent high cost and low reliability.
Some further problems encountered with the use of boosters generally include diiiiculty in separating the booster from the pay load. To provide some stability, lthe push type booster is connected at several circumferentially spacedpoints which must all be separated at once. Failure to achieve simultaneous separation at all points may contribute to altitude and velocity errors at the end of boost. Additionally, in the event the pay load itself constitutes ares or the like which are to be parachuted at a given altitude, the separated booster when at the rear of the pay load is in a position to become entangled with the chutes.
With all of the foregoing in mind, it is a primary object of the present invention to provide an improved booster rocket in which most of the above noted difticulties are overcome to the end that both direction and speed control for accurate ranging of ballistic missiles at the end of boost can be achieved.
More particularly, it is an object to provide an irnproved booster rocket in which any thrust ecceutricities 1 or misalignments are substantially eliminated and the effects of external atmospheric disturbances to the direc- 3,367,682 Patented Dec. 11, 1962 -booster rocket which is adaptable for connection to many different types of missiles and in which the connection itself need constitute only a single element to the end that separation for terminating the applied thrust is greatly facilitated and to the further end that once separated, the booster itself will not be in a position to interfere with any subsequent operation of the pay load.
Still another object is to provide a booster rocket in which substantially rectilinear flight paths for the pay load connected to the rocket can be realized with the resultant advantage that minimum hit ranges of line-ofsight controlled ground to ground missiles can be reduced to substantially zero.
Briefly, these as well as many further objects and vadvantages of this invention are attained by providing a booster rocket designed to pull, rather than push, the pay load. The pull rocket is positioned on the forward end of the missile and includes a thrust generating motor surrounded by a shroud structure preferably in the shape of a paraboloid or cone having an open end. An elongated stem co-inciding with the longitudinal axis of the pay load extends into the open end and supports the shroud `in axial alignment with the pay load. This stem constitutes the sole connection between the pull booster and pay load. The thrust generating motor may constitute for example, a series of rocket nozzles anranged in a circular array about the elongated stem Within the shroud butin spaced relationship thereto. A universal 'bearing rotatably mounts the nozzles to the elongated stem. For the case where the external conguration of the thrust motor is spherical or nearly spherical, that is, an external shape which essentially prevents thev existence of disturbing aerodynamic moments, the shroud itself may be eliminated.
By the `foregoing arrangement, the circular array of nozzles may be spun at gyroscopic speed so that the net thrust ydeveloped by the nozzles is stabilized in the direction of the axis of rotation of the nozzles. This high speed of rotation also cancels any thrust misalignments by averaging the same over 360 degrees. Because of the universal bearing mounting, the axis of the shroud stem and pay load can form an angle with the direction of thrust. The particular thrust direction however will always be constant as a consequence of the gyroscopic stability imparted thereto. Because the stem constitutes the sole connection of the pulling arrangement for towing the missile, release of the thrust generating system from the missile at the desired instant of time is greatly facilitated. 4Thus both direction and speed control of a pay load may be achieved by the gyroscopic pullrocket. In addition, the actual direction of thrust may be initially adjusted to provide a vertical thrust component substantially equal to the weight of the pull rocket and missile combined so that the entire system can be made to follow a horizontal path at a constant altitude above the ground.
A better understanding of the invention will be had by referring to one embodiment .thereof as illustrated in the accompanying drawings, in which:
FIGURE 1 is a perspective view of the pull rocket of ,this invention towing Aa pay load. v
FIGURE 2 is an enlarged cutJa-way perspective view of the pull rocket taken generally in the direction of the arrows 2--Z of FIGURE 1 showing one type of thrust generating system;
FIGURE 3 is a cross section of the universal bearing Afor the rocket motors Within the circular arrow 3 of FIGURE 2;
FIGURE 4 is another view similar to FIGURE 3 showing the bearing means in caged posi-tion; v `FIGURE 5 is a cut-a-way perspective view similar to TIGURE 2 showing a modified type of thrust generating neans;
FIGURE 6 shows one type of releasable connection :le-ans;
FIGURE 7 is lan elevational View of 4a pay load and he pull rocket of this invention, useful in explaining one if its applications; 'and FIGURE 8 is la diagrammatic view useful in explaining ome further advantages of the invention.
Referring first to FIGURE l there is illustrated -a typical ay load 10 which may be a ballistic missile incorporating warhead to which :an initial boost phase is to be imlarted. Secured to the nose portion of the pay load 10 t a pull rocket 11 including a shroud 12 surrounding ia trust gener-ating means 13.
In the particular illustration of FIGURE l, the longituinal axis of the pay load 10 and shroud 12 is designated i. The axis line T on the other hand represents the lrust taxis for the means 13 and it will be noted that this n'ust `axis 'forms an acute angle with the shroud Iaxis A.
Referring particularly to FIGURE 2, it will be evident iat lthe shroud 12 is exteriorly shaped generally lto corre- )ond to a paraboloid. This shape may vary however epend-ing upon `aerodynamic requirements in yaccordance 'ith whether or not the contemplated flight speed is submic or super-sonic. A single elongated stem 14 colciding with the axis of the paraboloid shape has one 1d secured -to the `interior no-se portion of the shroud s at 15 :and to the interior wall by ya suitable annular late or spider structure 16. The stem 14 is thus held in )Jaxial relationship with respect to the shroud. The ther end of the stem 14 terminates in a releasable coneotion 17 to the nose of the pay load 10.
The particular thrust lgenerating means 13 illustrated in IGURE 2, comprises :a series of rocket motors 18, 19, l), and 21, for example, each provided with rocket ozzles disposed in -a circular larray about the stem 14. hese motors are secured to a universal bearing housing Z by forward brackets 23 and rear brackets 24. The n iversal bearing housing 22 is shown in alignment with 1e thrust axis T `and this -axis lcorresponds to the taxis E the circular array of the various motors. When the lroud kand stem are in the dotted lline position,`the laxis co-incides with the `axis A. Y f
Each of the motors includes a small spin rocket such s indicated at 25 for the motor 18. These spin rockets re arranged to direct their thrust tangentially to the rcular array. The small spin rockets upon ning will lus cause the entire arr-ay to rotate with rthe universal caring housing 22 about an axis co-inciding with the xis T. 1
The universal bearing incorporated within the housing Z is shown in FIGURES 3 land 4 wherein the angle )rmed by the housing axis with the stem 14, in FIGURE is preserved in FIGURE 3. As shown, the interior '.alls of the housing 22 support -a bearing structure 26 lcluding upper and lower annular bearing races 27 fand 3, respectively. Co-operating with these bearing races re upper and lower lannular spherical bearing lsurfaces 29 nd 30 on the upper yand lower ends of a central core 31 idably surrounding the central portion of the stem 14. uitable upper `and lower ball bearings 32 and 33 comletethe bearing assembly. By this construction the stern 4 can elect universal movements within given limits 'ith respect to the bearing housing 22 to which the motors lemselves are rigidly secured. c
Caging and uncaging of the casing and bearing strucire, is achieved by making the central core 31, bearings., using, and motors secured thereto -all capable of limited ngitudinal movement as a unit along the central poron of the stem. Upward movement is limited by a )mpression spring 34 disposed between'the upper end E the core 31 and -a flange stop 35 txed to the stem. he lower end of the stem on the other hand includes a ging bear-ing 36 rigidly secured to the stem `and including an annular seat 37 receiving the lower end of the casing '22 when the structure is in caged position.
Thus, with particular reference to FIGURE 4, the stem and shroud axis is aligned with the thrust axis of the circular array of motors. Prior to launching of -a missile, the components are in the position illustrated in FIGURE 4, Iand held in caged position by the compression spring 34 and annular seat.
The operation of the caging mechanism is as follows: With the various component pants in .the position shown in FIGURE 4, the desired direction of thrust is determined and the axis 'of the missile, stem, and bearing housing 22 aligned in this direction.V This may be yaccomplished on 'the launcher by proper `adjustment' ofthe launching frame itself. Once the desired `direction of thrust has been established, the small spin rockets such as the rocket 25 of FIGURE 2 may be ignited to cause the circular `array of motors to rotate at high speed about the stem axis. The caging bearings 36 will maintain the axis of the bearing housing 22 and the axis of rota-tion of the motors in alignment with the stem kaxis so that the `universzd bearing surliaces 29 and 30 are locked against :any tilting but will permit the desired rotation to take place.
The rocket motors are spun to la speed sufcient to develop la gyroscopic action. When lthe main rocket motors are red, the Ithrust developed will overcome the force exerted by the compression spr-ing 34 `and the central bearing core 31 and housing will move upwardly along the stem from the position shown in FIGURE 4 to oompress the spring 34. The lower end of the housing will thus be free of the caging bearing as shown in FIGURE 3 so that the stem may then move in a universal sense Within the bearing housing. The uncaging of the motors lis accordingly yachieved automatically and in response to lthe development of thrust.
FIGURE 5 illustrates a modified thrust developing means which may be substituted for the motors 13 0f FIGURE 2. In FIGURE 5, the shroud, supporting stem, releasable connection, bearing means and caging assembly are identical to those shown in FIGURES 2, 3, and 4 and are designated by the same numerals.
VInstead of a plurality of separate motors secured to the bearing housing, however, there is provided asingle rocket motor casing 38 of generally spherical shape provided with individual rocket nozzles 39 disposed in a. circular array about the central axis of rotation of the bear ing housing and casing. Within the annular hollow interior of the casing 38 there are provided rings 40, 41, and 42 of solid propellant grain positioned between outer and inner annular plastic liners 43 and 44. By employ ing more than one ring, a larger 4burning area of the grain is exposed for more rapid combustion.
n Suitable spin-up motors 25 as in FIGURES 2, 3, and 4 may be employed to impart gyroscopic rotational speed to the casing 38. Ignition of the rings of propellant can be achieved by a centrifugal switch and exploding cartridge structure responsive to the rotation ofY the casing by the spin rockets and `secured within the casing itself as shown at 45.
In FIGURE 6 a preferred design of the releasably connection 17i`or the stem is shown in detail. A releasable connection is provided whenever the pull rocket is employed vfor purely ballistic type missiles for range control and the principal requirement is that the booster be released in response to a given signal at a given instance of time with minimum reaction on the pay load.
In the releasable connection 17 illustrated in IFIGURE 6, there is provided a coupling cavity 46 for receiving a projection 47 from the nose of the missile 10. The projection 47 includes an annular groove 48 receiving balls 49 partially projecting into the cavity 46.' An annular piston 50 in turn is positioned between the exterior of the cavity walls and interior of the connection 17 for sliding movement. Piston 50 includes annular sloping interior walls 51 defining ball receiving areas. A spring S2 urges the piston 50 in a rearward direction to lock the balls 49 in the groove 43 when the connection is attached. A small squib S is disposed in the interior of the rear portion of the projection and communicates with lateral ports P passing to the underside of the piston 50 when the connection is attached.
In operation, explosion of the squib S at the instant of time separation is desired is accomplished by an accelerometer on the missile or other transducer responsive to a physical condition of ight. Alternatively, the squib may be exploded from a ground command signal. The resulting expanding gases pass from the ports P to move the annular piston 50 upwardly against spring 52. The balls 49 can then yfall into the space 51 thus freeing the coupling cavity 46 from the projection 47.
The continued thrust of the booster will then simply pull the stem 14 and coupling 17 from the nose projection 47 of the missile with minimum reaction on the missile, leaving the missile free for ballistic flight. The payload will then follow a conventional ballistic trajectory to its target. Thus, critical launch angles need not be computed since the ballistic portion of the trajectory only extends from the moment of release of the pay load by the pull rocket.
The rapid spinning achieving the gyroscopic stability of the thrust `direction is one of the primary features of the instant invention and results in many useful applications of the pull rocket in addition to the foregoing described use with ballistic missiles. The ability to maintain a constant thrust direction can enable rectilinear flight paths of a missile to be achieved. For example, in FIGURE 7 there is shown a payload 53 being towed above the ground G by a pull rocket S4 such as described in FIGURES 2, 3, and 4. In this instance, the angle of thrust T with respect to the horizontal has been compute`d such that the thrust vector 55 provides a vertical component 56 exactly cancelling the weight of the payload 53 and pull rocket 54. The horizontal component 57 will provide the ydesired forward motion of the pay load so that the entire assembly will move in a horizontal path such as indicated by the line H.
The uses for straight line flight paths are many. As one example, a smoke screen curtain can be drawn over an extremely large area from a single pay load made to follow a rectilinear path at constant altitude. The stability alorded by the constant thrust direction will insure the desired straight line ight path.
FIGURE 8 illustrates another application of the pull rocket wherein the minimum hit range of a line of sight ground to ground -missile can be substantially reduced to zero.
In FIGURE 8, there is illustrated a ground to ground missile launcher 58 which would normally launch a missile along a conventional flight path indicated at 59 to strike a target 60. The path 59 represents the conventional path that the missile must follow in order to become properly airborne and before conventional aerody namic control thereof can be realized. It will be noted that the initial trajectory follows a parabolic path. There is thus la distance D within which the missile cannot be employed for kill purposes. The useful ranges of the missile therefore fall between the maximum range R and the minimum range D.
By providing a pull booster of the type described in FIGURE 7 on the missile launched at 58, however, a perfectly rectilinear flight path as shown by the dashed line 61 can be achieved throughout the entire range. Therefore, if a target were in the dotted line position 66 a direct hit could be scored. In fact, the entire path can be made at a constant altitude starting directly from the launcher so that the minimum distance heretofore decreasing the usefullness of ground to ground missiles can be substantially reduced to zero.
It will be understood of course that in addition to the principal feature of gyroscopic stabilization of thrust direction, the spinning of the thrust generating means achieves several other useful purposes. As already mentioned, thrust misalignments are cancelled as a consequence of the spinning operation. Further, in the event an extremely large pull rocket should be made employing liquid type fuel engines, the centrifugal forces established could be used for pumping the fuel thus eliminating the need for any fuel pumps.
By maintaining the axis of the shroud substantially in alignment with the pay load axis, the rotating rocket nozzles and motor are protected from adverse aerodynamic effects. Thus, the spinning structure is shielded from lateral aerodynamic forces which might cause procession and thus alter the direction of thrust. If a spherical type casing is employed as the thrust generating means such as shown in FIGURE 5, a shroud would not be an essential part of the structure.
Because of the fact that the pull rocket actually pulls the pay load, only a single connection at 17 is necessary. This feature renders the pull rocket adaptable to many different shaped missiles. Moreover, stability is achieved since the connection is in tension rather than compression and the pulling force is foreward of the center of gravity of the pay load. In addition, the continued thrust of the booster upon separation will cause it to rapidly pull away from the pay load once the connection has been released so that the booster can in no way interfere with subsequent operation of the pay load. Thus the pull rocket would be well suited for launching and dropping parachuted llares and the like.
From all of the foregoing, it will be seen that the present invention provides a greatly improved rocket booster. IWhile the word rocket has been employed throughout as describing the thrust generating means, the word as used in this connection is meant to include any type of thrust generating motor including liquid fuel engines, jets, and the like. Further, only two types of motors and nozzles have been shown as the thrust generating means, it should be understood that any annular type of construction which will provide thrust and enable rapid spinning of the thrust generating means about a central axis would be equivalent.
The invention is accordingly not to be thought of as limited to the particular embodiments chosen merely for illustrative purposes nor is it to be thought of as lim-` ited to the few applications set forth merely as exemplary.
What is claimed is:
l. A pull rocket comprising, in combination: a shroud structure generally in the shape of a paraboloid having an enlarged opening at its rear end; an elongated stem centrally positioned within said shroud and having one end portion coupled to said shroud, the other end of said stem terminating in connecting means adjacent to the open end of said shroud for connection to a pay load; thrust developing means positioned in an annular array about said stem within said shroud and in spaced relation to the interior wall of said shroud; and 4a universal bearing mounting said thrust developing means to said stern for rotation thereabout in such a manner that the axis of said stem can tilt with respect to the axis of rotation of said thrust developing means, whereby the thrust direction of s-aid thrust developing means can assume an acute angle with respect to the axis of said stem and pay load.
2. The subject matter of claim 1, in which said thrust developing means includes a series of rocket nozzles disposed in a circular array, the axis of said circular array defining the direction of net thrust developed by said nozzles; and spin rocket means secured to said thrust developing means to generate a thrust tangent to said circular array whereby said series of rocket nozzles can be spun at a speed about the axis of said circular array suflicient to provide gyroscopic stabilization of said direction of thrust.
3. The subject matter of claim 1, in which said conecting means between said stem and said pay load intudes a coupling structure having overlapping portions;
ad means responsive to a control signal -for moving said ortions out of overlapping relationship, whereby said ayload may be separated form said pull rocket lat a lven instant of time. l
4. The subject matter of claim 2, including caging leans for holding the axis of rotation of said thrust desloping means in co-axial alignment withv the axis of said em during spinning of said rocket nozzles, said caging leans including an annular caging bearing ixed to the wer portion of Saidr stern, said mounting means being idably mounted on said stem and including a lower poron receivable in4 said annular bearing when in its lowerlost position along said stem and when the axis of said em is co-axial with the axis of said mounting means, trust developed by said thrust developing means moving lid mounting means along said stem to a position free of ngagement with said caging bearing to uncage said lounting means form said stem.
5. The subject matter of claim 2, in which said thrust eveloping means includes la plurality of individual rocket rotors respectively connected to said rocketY nozzles dis :sed in an annular'array and connected to said mountlg means.k
6. The subject matter of clairn 2, Vin. which said thrust eyeloping means.v includes a single spherically shaped morr casing with a diametric opening incorporating said lounting means and through which said stem passes, lid nozzles being connected to exterior portions of said spherical casing to define said annular array and com.-
' municating with a single annular combustion chamber within said casing.
7. A pull rocket comprising: a supporting means adapted to be connected at one end to the nose portion of a pay load in axial alignment with the longitudinal axis of said pay load; a thrust developing means; mounting means for rotatably mounting said thrust developing means to said supporting means, said mounting means including a vuniversal bearing such that said supporting means and pay load may tilt with respect to the direction of thrust of said thrust developing meansgand means for rotating said thrust developing means to maintain ysaid direction of thrust constant by gyroscopic action.
8. The subject matter of claim 7, including a shroud structure surrounding said thrust developing means.
References Cited in the file of this patent UNITED STATES PATENTS 395,881 Cunningham Ian. 8, 1889 1,376,316 Chilowsky Apr. 26, 1921 1,748,697 Maier-Behrlng Feb. 25, 1930 2,145,508 Denoix Jan. 31, 1939 2,555,080 Goddard May 29, 1951 12,792,758 Bach et al. May 21, 1957 2,849,955 Smathers Sept. 2, 1958 2,879,668 Mleczko Mar. 3l, 1959 2,884,859 Alexander et al. May 5, 1959 2,945,442 Adelman et al. n July 19, 1960 3,001,739 Faget et al. Sept. 26, 1961