GB2121147A - Missile fin assemblies - Google Patents
Missile fin assemblies Download PDFInfo
- Publication number
- GB2121147A GB2121147A GB08315131A GB8315131A GB2121147A GB 2121147 A GB2121147 A GB 2121147A GB 08315131 A GB08315131 A GB 08315131A GB 8315131 A GB8315131 A GB 8315131A GB 2121147 A GB2121147 A GB 2121147A
- Authority
- GB
- United Kingdom
- Prior art keywords
- fin member
- missile
- fin
- link
- link means
- 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.)
- Granted
Links
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
- F42B10/16—Wrap-around fins
Landscapes
- Physics & Mathematics (AREA)
- Fluid Mechanics (AREA)
- Engineering & Computer Science (AREA)
- General Engineering & Computer Science (AREA)
- Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)
Abstract
A missile includes a fin member 1 hingedly mounted for hinging movement about hinge axis H between a stowed position and a deployed position. The fin member 1 includes a transverse lever arm 4 which pivotally carries one end of a link 5. The other end of the link 5 is pivotally coupled to a cylindrical shell 3, which is relatively rotatable with respect to hinge axis H about a rotation axis R to move the other end of the link along a circular path thus to move the fin member to its deployed position. In a preferred arrangement, when the fin member is in its deployed position, the lever arm 4 extends tangentially to axis R and link 5 extends radially with respect thereto to provide a geometric lock. <IMAGE>
Description
SPECIFICATION
Missile fin assemblies
This invention relates to missiles which include one or more fin members movable between a stowed and deployed condition.
According to this invention, there is provided a missile including a fin member, link means and drive means, said fin member being hingedly mounted for hinging movement about a generally longitudinal hinge axis from a stowed position in which it lies generally alongside a region of the missile body and a deployed position in which it projects therefrom, said fin member including lever arm means extending generally transversely of said hinge axis, said link means being pivotally coupled at one end to said arm means and having its other end region constrained to move along a generally circular arc centred on a rotation axis, said drive means being arranged to move said other end region of the link means along the generally circular arc to effect movement of the fin member from its stowed position to its deployed position,
The fin member and the link means are preferably so configured that when the fin member is in its deployed position, said link means is aligned substantially radially with respect to said rotation axis and said lever arm means is aligned substantially tangentially with respect thereto. This deployed configuration provides a geometric lock against movement of the deployed fin member and, furthermore, ensures that for a given rate of movement of said other end of the link means, the rate of deployment of the fin member decreases as it approaches its deployed position.This is because the lever arm means and the link means are in a dead centre position when the fin member is in its deployed position.
Conveniently, said fin member is hingedly attached to an outer shell member and said drive means includes
an inner shell member arranged concentrically within said outer shell member, the inner and outer shell members being relatively rotatable with respect to said rotation axis.
Preferably, abutment means are provided to restrain movement of said link means when the fin member is in its deployed position. The missile preferably is provided with a piurality of fin members around its periphery and each fin member is provided with associated link means, each of these link means being coupled to a common drive member.
By way of example only, certain specific embodiments of missile will now be described in detail, reference being made to the accompanying drawings, in which:
Figure 1 a, 1 b and 1 c are transverse sectional views of the tail unit of a missile and show a fin assembly in stowed, intermediate and deployed conditions respectively;
Figure 2 is a longitudinal sectional view of part of a tail unit of a missile showing a fin assembly in its deployed condition; and
Figure 3 is a transverse sectional view through a further form of tail unit for a missile.
Referring initially to Figure 1, this shows a single fin member 1 which is hingedly attached to outer cylindrical shell 2 for pivotting movement about hinge axis H for movement between a stowed configuration (Figure 1 a) and a deployed condition (Figure 1 c). The fin member 1 is curved to lie ciosely adjacent shell 2 when stowed. The
cylindrical shell 2 is fixed to, or forms an extension of, the missile body. An inner drive shell 3 is
arranged coaxially within shell 2 for turning
movement about rotation axis R. The fin member
1 includes a lever arm 4 extending transversely of
hinge axis H and which is pivotally coupled to a point on the inner shell 3 by means of link 5. The
drive shell 3 is biassed in the clockwise direction
by means of a spring device (not shown) and is
latched in the stowed configuration by latch
means (not shown).On release of the latch means, the drive shell is urged clockwise thereby causing the arrangement to pass through the
intermediate configuration (Figure 1 b) to the
deployed condition (Figure 1 c). The arrangement
is configurated so that when in the deployed
condition, link 5 is aligned radially with respect to
rotation axis R, abutting abutment region 6 of the
cylindrical shell 2. The radial alignment provides a
geometric lock to prevent inertial or aerodynamic
loads on the fin member 1 causing hinging
movement thereof. Link 5 is further prevented
from angular movement with respect to the
drive shell 3 in on sense by abutment region 6 and
in the other sense by the bias of the spring.
Moreover, the lever arm is configured so that
when the fin member is erected, it lies
substantially perpendicular to link 5; this
configuration means that any turning moment
imparted to the fin member 1 when erected
results in a substantially wholly axial load on link
5. As a yet further precaution against movement
of the erected fin member, the drive shell includes
an abutment region 7 so that when in the erected
condition link 5 is clamped between abutment
regions 6 and 7.
Turning to Figure 2, each fin member is
provided with three lever arms 4 arranged in
tandem along the hinge axis H, each lever arm
being coupled to the drive shell 3 by means of an
associated link 5. Three lever arms 4 are preferred
in the case shown, but other cases could involve
more or less arms.
The arrangement of Figure 2 also differs from
that of Figure 1 in that outer cylindrical shell 2 is
rotatably mounted with respect to the missile
body whilst inner shell 3 is fixed with respect
thereto. In this case, erection and stowage
movement will be imparted to the outer shell 2.
In another case, both outer and inner shells 2
and 3 may be free from constraint with respect to
the missile body. In this latter case, inner shell 3
must be located by outer shell 2 in rotational
relationship therewith, bearings being provided as
shown between the missile body and shell 2 or
between the body and shell 3.
Referring now to Figure 3, a typical arrangement employing four fin members is shown. In this arrangement each fin member 1 is moved in the same manner as in the arrangements of Figures 1 or 2 via a common inner drive shell 3.
In this arrangement, any transient load imparted to one of the fin members will be transmitted to the other fin member, so that transient forces acting on different fins at the same time will be balanced to some extent by the opposing forces transmitted to the drive shell 3.
Furthermore, for a given rate of rotation of shell 3 relative to shell 2, the rate of hinging of the fins to their deployed positions decreases as link 5 approaches its radially aligned position with respect to rotation axis R. The deployment motion of the fin is therefore arrested with reduced shock load on any part of the mechanism compared to previous arrangements in which the fin member itself is urged by a spring to meet the abutment stop when the fin member reaches its deployed position. Consequently, the internal forces acting during fin erection and locking may be reduced and regulated, thereby reducing the loads transmitted through the hinge and locking arrangement of the fin assembly. This may enable lighter and more compact arrangements to be designed.
Claims (6)
1. A missile including a fin member, link means and drive means, said fin member being hingedly mounted for hinging movement about a generally longitudinal hinge axis from a stowed position in which it lies generally alongside a region of the
missile body and a deployed position in which it
projects therefrom, said fin member including lever arm means extending generally transversely of said hinge axis, said link means being pivotally coupled at one end to said arm means and having its outer end region constrained to move along a generally circular arc centred on a rotation axis, said drive means being arranged to move said other end region of the link means along the generally circular arc to effect movement of the fir' member from its stowed position to its deployed position.
2. A missile according to Claim 1, wherein the fin member and the link means are so configured that when the fin member is in its deployed position, said link means is aligned substantially radially with respect to said rotation axis and said lever arm means is aligned substantially tangentially with respect thereto.
3. A missile according to Claim 1 or Claim 2, wherein said fin member is hingedly attached to an outer shell member, and said drive means includes an inner shell member arranged concentrically within said outer shell member, the inner and outer shell members being relatively rotatable with respect to said rotation axis.
4. A missile according to any of the preceding
Claims, wherein abutment means are provided to at least restrain movement of said link means when the fin member is in its deployed position.
5. A missile according to any of the preceding
Claims, wherein a plurality of fin members are provided around the periphery of the missile, each fin member having associated link means and each link means being coupled to a common drive member.
6. A missile substantially as hereinbefore described with reference and as illustrated in any of the accompanying drawings.
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
GB08315131A GB2121147B (en) | 1982-06-02 | 1983-06-02 | Missile fin assemblies |
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
GB8215895 | 1982-06-02 | ||
GB08315131A GB2121147B (en) | 1982-06-02 | 1983-06-02 | Missile fin assemblies |
Publications (2)
Publication Number | Publication Date |
---|---|
GB2121147A true GB2121147A (en) | 1983-12-14 |
GB2121147B GB2121147B (en) | 1985-10-16 |
Family
ID=26283002
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
GB08315131A Expired GB2121147B (en) | 1982-06-02 | 1983-06-02 | Missile fin assemblies |
Country Status (1)
Country | Link |
---|---|
GB (1) | GB2121147B (en) |
Cited By (11)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP0157112A1 (en) * | 1984-03-09 | 1985-10-09 | Rheinmetall GmbH | Fin-stabilised projectile |
FR2571133A1 (en) * | 1984-10-01 | 1986-04-04 | Commw Of Australia | WING DEPLOYMENT MECHANISM, IN PARTICULAR FOR MISSILE |
EP0242180A2 (en) * | 1986-04-15 | 1987-10-21 | British Aerospace Public Limited Company | Deployment arrangement for spinning body |
FR2617587A1 (en) * | 1987-07-03 | 1989-01-06 | Thomson Brandt Armements | DEVICE FOR CONJUGATED DEPLOYMENT OF WINGS, AND APPLICATION TO A FLYING DEVICE |
US4844381A (en) * | 1987-09-08 | 1989-07-04 | Diehl Gmbh & Co. | Airborne submunition member |
US5031856A (en) * | 1989-05-12 | 1991-07-16 | Diehl Gmbh & Co. | Airborne submunition member |
US5816532A (en) * | 1996-12-17 | 1998-10-06 | Northrop Grumman Corporation | Multiposition folding control surface for improved launch stability in missiles |
RU2520846C1 (en) * | 2013-03-29 | 2014-06-27 | Открытое акционерное общество "Военно-промышленная корпорация "Научно-производственное объединение машиностроения" | Rocket aerodynamic rudder |
RU2587751C1 (en) * | 2015-03-16 | 2016-06-20 | Акционерное общество "Военно-промышленная корпорация "Научно-производственное объединение машиностроения" | Deployable rudder |
EP3111157A4 (en) * | 2014-02-26 | 2017-09-27 | Israel Aerospace Industries Ltd. | Fin deployment system |
US10323917B2 (en) | 2013-10-10 | 2019-06-18 | Bae Systems Bofors Ab | Fin deployment mechanism for projectile and method for fin deployment |
-
1983
- 1983-06-02 GB GB08315131A patent/GB2121147B/en not_active Expired
Cited By (14)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP0157112A1 (en) * | 1984-03-09 | 1985-10-09 | Rheinmetall GmbH | Fin-stabilised projectile |
FR2571133A1 (en) * | 1984-10-01 | 1986-04-04 | Commw Of Australia | WING DEPLOYMENT MECHANISM, IN PARTICULAR FOR MISSILE |
EP0242180A2 (en) * | 1986-04-15 | 1987-10-21 | British Aerospace Public Limited Company | Deployment arrangement for spinning body |
EP0242180A3 (en) * | 1986-04-15 | 1989-04-26 | British Aerospace Public Limited Company | Deployment arrangement for spinning body |
FR2617587A1 (en) * | 1987-07-03 | 1989-01-06 | Thomson Brandt Armements | DEVICE FOR CONJUGATED DEPLOYMENT OF WINGS, AND APPLICATION TO A FLYING DEVICE |
EP0298844A1 (en) * | 1987-07-03 | 1989-01-11 | Thomson-Brandt Armements | Device for deploying fins and its use in a projectile |
US4844381A (en) * | 1987-09-08 | 1989-07-04 | Diehl Gmbh & Co. | Airborne submunition member |
US5031856A (en) * | 1989-05-12 | 1991-07-16 | Diehl Gmbh & Co. | Airborne submunition member |
US5816532A (en) * | 1996-12-17 | 1998-10-06 | Northrop Grumman Corporation | Multiposition folding control surface for improved launch stability in missiles |
RU2520846C1 (en) * | 2013-03-29 | 2014-06-27 | Открытое акционерное общество "Военно-промышленная корпорация "Научно-производственное объединение машиностроения" | Rocket aerodynamic rudder |
US10323917B2 (en) | 2013-10-10 | 2019-06-18 | Bae Systems Bofors Ab | Fin deployment mechanism for projectile and method for fin deployment |
EP3111157A4 (en) * | 2014-02-26 | 2017-09-27 | Israel Aerospace Industries Ltd. | Fin deployment system |
US9989338B2 (en) | 2014-02-26 | 2018-06-05 | Israel Aerospace Industries Ltd. | Fin deployment system |
RU2587751C1 (en) * | 2015-03-16 | 2016-06-20 | Акционерное общество "Военно-промышленная корпорация "Научно-производственное объединение машиностроения" | Deployable rudder |
Also Published As
Publication number | Publication date |
---|---|
GB2121147B (en) | 1985-10-16 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
US4296895A (en) | Fin erection mechanism | |
US4381092A (en) | Magnetic docking probe for soft docking of space vehicles | |
GB2121147A (en) | Missile fin assemblies | |
EP1550837B1 (en) | Device for the deployment and the control of the control vanes of a projectile | |
US6010096A (en) | Deployment restraint and sequencing device | |
US4564160A (en) | Thrust reverser blocker door assembly | |
EP0704373B1 (en) | Deployment hinge apparatus | |
US5257034A (en) | Collapsible apparatus for forming a paraboloid surface | |
EP1013546B1 (en) | Rocket payload fairing and method for opening same | |
EP1485668B1 (en) | Deployment mechanism for stowable fins in missiles | |
EP1386838B1 (en) | Deployable antenna reflector | |
US4880188A (en) | Joint for unfolding panels of a solar collector | |
GB2150092A (en) | Deployment and actuation mechanisms | |
US4284387A (en) | Blade fold restraint system | |
US6199988B1 (en) | Retractable device, of the light shield type, for an optical instrument such as a space telescope | |
EP1783051A1 (en) | Combination actuator latch mechanism | |
JP2755492B2 (en) | Missile with deployable steering wings | |
AU2020407421A1 (en) | Device for damping docking to a satellite | |
US4350297A (en) | Swivelling exhaust nozzles for rocket motors | |
US5615632A (en) | Underwater vehicle and a fin assembly therefor | |
US5816532A (en) | Multiposition folding control surface for improved launch stability in missiles | |
US3724782A (en) | Deployable aerodynamic ring stabilizer | |
US6308919B1 (en) | Spacecraft having a dual reflector holddown for deploying multiple reflectors in a single release event | |
GB2369420A (en) | Device for exerting drag on a projectile in flight | |
US5738308A (en) | Ion thruster support and positioning system |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
PCNP | Patent ceased through non-payment of renewal fee |