US3319566A - Non-spin rockets and their guidance - Google Patents

Non-spin rockets and their guidance Download PDF

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Publication number
US3319566A
US3319566A US570124A US57012466A US3319566A US 3319566 A US3319566 A US 3319566A US 570124 A US570124 A US 570124A US 57012466 A US57012466 A US 57012466A US 3319566 A US3319566 A US 3319566A
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US
United States
Prior art keywords
rocket
nozzle
motor
charge
gas stream
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
US570124A
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English (en)
Inventor
Paul V Choate
Michael A Nee
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Norris Industries Pty Ltd
Norris Industries Inc
Original Assignee
Norris Industries Pty Ltd
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Norris Industries Pty Ltd filed Critical Norris Industries Pty Ltd
Priority to US570124A priority Critical patent/US3319566A/en
Application granted granted Critical
Publication of US3319566A publication Critical patent/US3319566A/en
Priority to NO168189A priority patent/NO122905B/no
Priority to SE9747/67A priority patent/SE333317B/sv
Priority to FR115517A priority patent/FR1532158A/fr
Priority to BE701786D priority patent/BE701786A/xx
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B10/00—Means for influencing, e.g. improving, the aerodynamic properties of projectiles or missiles; Arrangements on projectiles or missiles for stabilising, steering, range-reducing, range-increasing or fall-retarding
    • F42B10/02—Stabilising arrangements
    • F42B10/14—Stabilising arrangements using fins spread or deployed after launch, e.g. after leaving the barrel
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B10/00—Means for influencing, e.g. improving, the aerodynamic properties of projectiles or missiles; Arrangements on projectiles or missiles for stabilising, steering, range-reducing, range-increasing or fall-retarding
    • F42B10/02—Stabilising arrangements
    • F42B10/26—Stabilising arrangements using spin
    • F42B10/28—Stabilising arrangements using spin induced by gas action
    • F42B10/30—Stabilising arrangements using spin induced by gas action using rocket motor nozzles

Definitions

  • FIG. 1 A first figure.
  • the present invention relates to non-spin rockets and to their guidance, the rockets being of any size ranging from those that are pad-based to small rockets such as .those that may be launched from shoulder-supported launchers.
  • the present invention is a continuation-inpart. of our copending application, Ser. No. 446,275, filed Apr. 7, 1965, and now abandoned.
  • Rockets like other projectiles, require stabilization and, except for the present invention and apart from guidance systems, there are two known methods by which stabilization can be eflected.
  • the two methods as set forth in ORDP 20-246, Ordnance Corporation Pamphlet, Ordnance Engineering Design Handbook, Artillery Ammunition Series, Section 3, Design for Control of Flight Characteristics, are spin stabilization and fin stabilization. If a projectile is spin stabilized, any increase in its length increases the difliculty of or makes it impossible to effect its stabilization. For that reason, there is a diameterto-length relationship that is characteristic of spin rockets,
  • a primary objective of the present invention is, accordingly, to provide rockets of increased stability in flight by a third basis of stabilization and this objective is attained by providing a rocket having a motor with the rocket having an elongate-d, rearwardly extending portion and the motor having a rearwardly opening nozzle surrounding the rocket approximately in the zone of the center of gravity of the rocket and the rearward rocket portion being dimensioned to be enveloped by the expanding gas stream from the nozzle and, being free of surface portions that would so interrupt the gas stream as to afiect adversely its holding action as would be the case if the rocket were spun.
  • a rockets accuracy is also affected by its velocity as is its range.
  • Another primary objective of the invention is, accordingly, to provide high velocity rockets even when the rockets are to be fired from shouldersupported launchers.
  • This objective is attained by providing for a rocket in accordance with the invention being launched with a propellant charge, sometimes herein referred to as the flight assisting or sustaining charge, effective when the rocket is air borne to provide the desired velocity, such charge being ignited either within the launcher or when air borne or becoming fully effective when air borne, the rocket being launched at a relatively low velocity as by a charge, sometimes herein referred to as the launching charge, contained in the rocket or within the launcher, or by providing that the flight assisting or sustaining charge be not fully effective until the rocket is air borne.
  • a propellant charge sometimes herein referred to as the flight assisting or sustaining charge
  • the opportunity for eccentric gas stream forces to affect a rockets course is not only minimized by the manner in Which the accelerating forces are applied thereto but also such forces are prevented from so doing by the use of the expanding gas stream as a gas envelope holding the rocket on course with a long moment arm advantage.
  • tail fins are desirable when the fuel supply burns so quickly that the thrust is applied to the rocket during only a minor part of its flight.
  • the motor is so designed as to provide thrust throughout a predetermined range of a rocket, fins are not necessary.
  • rotation of that rocket is desirable only as a means for increasing its accuracy by offsetting the efiect of any unevenness in the gas flow through the annular nozzle.
  • Another objective of the invention is, accordingly, to provide a rocket that rotates in flight at a rate far below a rate useful in spin stabilization thus to avoid upsetting the holding action of the gas stream, the rotative forces being applied thereto in the zone of its center of gravity.
  • the gas stream may be used to effect changes in course in conjunction with various guidance systems such as adjustable tail fins, longitudinal fins in the gas stream that can be turned, thrust deflectors in the nozzle, or by the use of radial gas jets, preferably in the area of the gas envelope.
  • various guidance systems such as adjustable tail fins, longitudinal fins in the gas stream that can be turned, thrust deflectors in the nozzle, or by the use of radial gas jets, preferably in the area of the gas envelope.
  • FIGURE 1 is a longitudinal section of a rocket in accordance with the invention, the rocket being of a type utilizing both launching and booster or flight sustaining propellant charges,
  • FIGURE 2 is a fragmentary view of the throat of another motor in accordance with the invention, the throat being defined in part, by an internal plug,
  • FIGURE 3 is a like view of the throat of another embodiment of the invention illustrating an adjustable internal plug
  • FIGURE 4 is a somewhat schematic view of the rocket of FIGURE 1 in flight, illustrating the pressure envelope provided by the gas stream,
  • FIGURE 5 is a fragmentary and partly sectioned view illustrating the annular nozzle of another rocket in accordance with the invention
  • FIGURE 6 is a section transversely of the rocket illustrated by FIGURE 5 rearward-1y of its nozzle
  • FIGURE 7 is a like view illustrating a rocket of the type illustrated by FIGURES l-4 and its launcher
  • FIGURE 8 is a similar view illustrating another rocket in accordance with the invention, the rocket also being of a type utilizing both launching and booster or flight sustaining propellant charges, the launching charge being of the external motor type,
  • FIGURE 9 is a like view of yet another similar rocket in accordance with the invention and its launcher of the recoilless rifle type,
  • FIGURE 10 is a similar view of another rocket in accordance with the invention and its launcher, the rocket having a single charge, and
  • FIGURE 11 is a similar view of the rocket whose nozzle is illustrated in FIGURES 5 and 6, the rocket also having a single charge and capable of being launched from a launcher such as that illustrated by FIGURE 10.
  • the rocket detailed in FIGURE 1 includes a launching motor, a flight sustaining or booster motor, and a war head generally indicated at 10, 11, and 12, respectively.
  • the launching motor 10 consists of a tubular chamber 13 for a fuel supply 14 provided with an igniter 15 in an indicated ignition circuit 16.
  • the launching motor 10 also has a rearwardly opening, axial nozzle 17.
  • the out side diameter of the chamber 13 is increased as at 18 at the forward end and is interiorly threaded to receive a closure 19.
  • the closure 19 is shown as having a central socket 20 threaded to receive the axial projection 21 of a cap 22.
  • Fins 23 are connected by pivots 24 to ears 25 on the nozzle 17 so that they may be folded forwardly into their dotted line positions shown in FIGURE 1 when the rocket is inserted into its launcher 26, see FIGURE 7.
  • the maximum diameter of the nozzle 17 may be greater than the diameter of the motor 10 being limited only to the bore dimension of the launcher.
  • Springs 27, connected to the fins 23 and to the ears 25, are operative to urge the fins 23 into their erected position, once the rocket is air borne.
  • the flight sustaining motor 11 consists of a cylindrical, rearwardly opening housing 28 having an axial bore 29 in its front wall 30 threaded to receive the forward end of a post 31 the rear end of which is threaded into a socket 32 in the front face of the cap 22.
  • the post 31 has a bore 33 opening through its front end and a series of radial ports 34 effecting communication between the bore 33' and the interior of the housing 28.
  • the fuel supply of the motor 11 is shown as in the form of a sleeve 35 located within the housing 28 and within the bore 33 there is an igniter 36 in an indicated ignition circuit 37.
  • the inner surface of the housing 28, at the rear end thereof, is outwardly flared as at 38 and the housing 28 is so dimensioned relative to the front part of the chamber 13 that the annular flared surface 38 overlies the junction of the chamber portion 18 and the rounded edge of the cap 22 to define a generally indicated annular nozzle 39. It is essential to note that the nozzle 39 is approximately in the zone of the center of gravity of the rocket as will be apparent from FIGURES 4 and 7.
  • the front face of the housing wall 30 has an annulus 40 in support of a sleeve 41 receiving the inner end 42 of the war head 43, the front face 44 of which supports an axial, collapsible nose 54.
  • a faring 46 of the same outside diameter as the housing 28, completes the front part of the rocket.
  • the rocket just disclosed is placed in its launcher 26.
  • the circuit 16 is closed and the launching motor 10 becomes operative.
  • the motor 11 does not become operative until the rocket is air borne, an effect attained as by providing, for example, that the circuit 37 be closed with the closing of the circuit 16 but that the igniter 36 be relatively slow burning.
  • the motor 10, accordingly, is effective to place the rocket in flight but the principal thrust is derived from the motor 11. Gas from the motor 11 is delivered through the rearwardly disposed, annular nozzle 39 with the thrust being applied annularly and approximately at the center of gravity of the rocket.
  • the propelling stream of expanding gas is in the form of an envelope 47 applying increased pressure against the rear part of the nozzle so that while the rocket is subjected to propelling forces, it is held on course by the gas stream, ensuring both accuracy and long range.
  • FIGURE 2 in which the motor 10A has its throat defined in part by an internal plug 48 underlying the end 38A of the housing 28A of the motor 11A and also to FIGURE 3 wherein the internal plug 48A is in engagement with threads 49 on the motor 10B to enable its position to be varied relative to the shoulder 38B of the housing 28B of the motor 113.
  • Rockets of other types may be made in accordance with the invention.
  • the rocket shown in FIG- URE 8 is adapted to be fired from the launcher 50, and has a launching motor, a flight sustaining motor, and a war head generally indicated at 10C, 11C, and 12C, respectively.
  • the war head 12C may be identical to the war head 12 and its motor is identical to the motor 11 except that, in place of the post 31, there is an axial rod 51 having a member 52 defining with the housing 28C the annular nozzle 39C approximately in the zone of the center of gravity of the rocket when in flight.
  • the rod has a second member 53 at its rear end adjacent which are fins 54.
  • the fuel supply of the motor NC is shown in dotted lines and is indicated at 14C, the dotted line showing being necessary because the rocket of FIG- URE 8 is illustrated as air borne.
  • the launcher serves as the casing of the motor 10C and defines, with the second member 53, its nozzle.
  • the rocket of FIGURE 9 is for use with the launcher 55 and, like the rocket of FIGURE 8, its launching motor 10D is completed by its launcher 55 when loaded therein.
  • the launcher 55 has a nozzle 56 at its rear end, the rod member 51D, while having a member 52D defining an annular nozzle 39D with the housing 28D of the motor 11D, is not provided with a member such as the member 53 but the rocket of FIGURE 9 may otherwise be identical to the rocket of FIGURE 8 and its annular nozzle 39D is approximately in the zone of its center of gravity.
  • the rocket illustrated by FIGURE 10 like that shown in the co-pending application of Paul V. Choate, Ser. No. 395,242, filed Sept. 9, 1964, is to be launched from the launcher 57 and has a single motor 58 and a war head 12E, the single motor having its annular nozzle 39E approximately in the zone of the center of gravity of the rocket.
  • this rocket can be launched from the launcher 57, has a single motor 59 having its annular nozzle 39F approximately in the zone of the rockets center of gravity and includes a war head 12F.
  • the motor 59 is adapted to provide thrust throughout the maximum range of the rocket and the rocket does not have tail fins.
  • the nozzle 39F is shown as having three fins 60 spaced apart in its throat and disposed angularly with respect to the ro-ckets axis whereby the rocket is rotated by the gas stream through the nozzle 39F, the rate of rotation being desirably in the vicinity of 15 to 25 r.p.s.
  • a motor having a chamber, and a propellant charge in said chamber, said motor including a rearwardly opening annular nozzle having a throat and surrounding said rocket in a zone approximately inclusive of the center of gravity of said rocket during the burning of said charge, said motor including a mixing chamber having rearward, direct, annular and substantially unobstructed communication with the throat of said nozzle, a line lengthwise of any portion of the part of said rocket rearwardly of said nozzle being capable of being presented to the viewer as a fore-and-aft line of the rocket at least when the rocket is in flight, the part of said rocket rearwardly of said nozzle being cylindrical and of a diameter less than the outside diameter of said nozzle in a zone commencing immediately adjacent the nozzle and extending substantially throughout its length, and the nozzle being disposed to direct the propelling gas stream resulting from the burning of said charge substantially parallel to the flight axis of the rocket and close to said rearward part along the entire cylindrical length thereof whereby, during the burning of
  • the rocket of claim 1 in which the portion of the rocket rearwardly of the nozzle includes a second motor having a chamber, a propellant in the chamber of the second mot-or, and the second motor includes a rearwardly.
  • each portion has a front end wall and an axial post interconnects the end walls, the rear end of the forward portion being open and overlying the forward end of the rearward cylindrical portion and establishing therewith the rearwardly opening, annular nozzle.

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  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Engineering & Computer Science (AREA)
  • General Engineering & Computer Science (AREA)
  • Toys (AREA)
US570124A 1966-07-25 1966-07-25 Non-spin rockets and their guidance Expired - Lifetime US3319566A (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
US570124A US3319566A (en) 1966-07-25 1966-07-25 Non-spin rockets and their guidance
NO168189A NO122905B (sv) 1966-07-25 1967-05-18
SE9747/67A SE333317B (sv) 1966-07-25 1967-06-29 Anordning vid icke rotationsstyrd raket
FR115517A FR1532158A (fr) 1966-07-25 1967-07-25 Nouveau procédé de stabilisation non gyroscopique des fusées engins
BE701786D BE701786A (sv) 1966-07-25 1967-07-25

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US570124A US3319566A (en) 1966-07-25 1966-07-25 Non-spin rockets and their guidance

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US3319566A true US3319566A (en) 1967-05-16

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US570124A Expired - Lifetime US3319566A (en) 1966-07-25 1966-07-25 Non-spin rockets and their guidance

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BE (1) BE701786A (sv)
NO (1) NO122905B (sv)
SE (1) SE333317B (sv)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3390850A (en) * 1967-08-04 1968-07-02 Army Usa Fin for inducing spin in rotating rockets
US4046076A (en) * 1975-09-29 1977-09-06 The United States Of America As Represented By The Secretary Of The Navy Impulsive rocket motor safety-arming device
US5596166A (en) * 1994-12-28 1997-01-21 Logicon Rda Penetrating vehicle with rocket motor

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2503271A (en) * 1945-02-06 1950-04-11 Clarence N Hickman Rocket projectile
US2946261A (en) * 1956-05-02 1960-07-26 Sydney R Crockett Peripheral nozzle spinner rocket
US3176615A (en) * 1962-12-31 1965-04-06 Avco Corp Gun-propelled rocket-boosted missile

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2503271A (en) * 1945-02-06 1950-04-11 Clarence N Hickman Rocket projectile
US2946261A (en) * 1956-05-02 1960-07-26 Sydney R Crockett Peripheral nozzle spinner rocket
US3176615A (en) * 1962-12-31 1965-04-06 Avco Corp Gun-propelled rocket-boosted missile

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3390850A (en) * 1967-08-04 1968-07-02 Army Usa Fin for inducing spin in rotating rockets
US4046076A (en) * 1975-09-29 1977-09-06 The United States Of America As Represented By The Secretary Of The Navy Impulsive rocket motor safety-arming device
US5596166A (en) * 1994-12-28 1997-01-21 Logicon Rda Penetrating vehicle with rocket motor

Also Published As

Publication number Publication date
BE701786A (sv) 1968-01-25
SE333317B (sv) 1971-03-08
NO122905B (sv) 1971-08-30

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