GB2164612A - Vehicles fitted with thrust vector control systems - Google Patents

Abstract

A vehicle, e.g. a guided missile, fitted with thrust vector control apparatus has a plurality of exhaust nozzles (4, 5) mounted for universal movement about respective pivot points (6), aerodynamic fins (12) mounted for pivotal movement about respective pivot axes, and a steering control system comprising actuators (7) (only one shown) operable to effect pivotal movement of the nozzles (4, 5), with link arms (14) extending between the nozzles and the said fins, the arrangement being such that each nozzle (4, 5) is coupled to at least two of said fins (12) and pivotal movement of each nozzle results in movement of its associated fins about their respective pivot axes. Link arms (14) connect the nozzles (4, 5) to crank arms (9), and pistons (8) of the actuators (7) are each connected to one end of a crank arm (9), the other end of which is attached to a shaft (11) on which an aerodynamic fin (12) is mounted. The vehicle may, alternatively, be a watercraft, e.g. an underwater vehicle, in which case the fins provide hydrodynamic control. <IMAGE>

Description

SPECIFICATION Vehicles fitted with a thrust vector control system This invention relates to vehicles fitted with thrust vector control systems and more particularly to steering control systems for such vehicles. In the main, the invention will be discussed in relation to guided missiles but it is to be understood that it is not restricted thereto and may, for example, be applied to watercraft, especially underwater vehicles.

In the early days of guided missiles, guidance or steering thereof was effected by adjustable aerodynamic surfaces the angular orientation of which was controlled by actuators to produce the desired missile manoeuvre, the necessary forward velocity being provided by a motor giving a undirectional thrust. With the advent of a thrust vector control system for a guided missile, the flight control function was transferred to that system and the adjustable aerodynamic surfaces dispensed with. Such a missile has proved satisfactory because the thrust vector control system makes the missile highly manoeurvrable but it has one limitation in that once the motor becomes inoperative, no further control of its flight can be effected.

The present invention is based on the realisation that a more versatile missile would result from a steering or guidance control system which combines thrust vector control with aerodynamic control. Indeed a dual purpose missile would result in that it would be highly manoeuvrable during the boost phase of the flight, which is normally immediately after launch when the motor is producing maximum thrust, and thus ideally suited to dogfight situations and/or to launching in any direction relative to a target. Following motor burnt-out, the missile can be steered in the glide phase by adjustable fins or other controllable aerodynamic surfaces, as long as sufficient forward velocity is maintained and this gives a second role or purpose to the missile, namely that of seeking out a long range target.A missile may have a coast phase between the boost and glide phases in which the motor produces a reduced thrust but still sufficient for thrust vector control to be implemented.

According to the present invention there is provided a vehicle fitted with thrust vector control apparatus having a plurality of nozzles mounted for universal movement about respective pivot points, aerodynamic or hydrodynamic surfaces mounted for pivotal movement about respective pivot axes, and a steering control system comprising actuator means operable to effect pivotal movement of the nozzles, and linkage means extending between the nozzles and said surfaces, the arrangement being such that each nozzle is coupled to at least two of said surfaces and pivotal movement of the nozzle results in movement of the associated surfaces about their respective pivot axes.

The aerodynamic or hydrodynamic surfaces may be in the form of tail fins which are thus mounted on the body of the missile in the vicinity of the thrust vector control nozzles. The angular relationship between the pivot axes of the fins and the pivot point of the associated nozzle is optional but in the interests of simplified linkages and of geometric relationships, it is preferable to arrange the pivot axis of each fin such that it subtends an angle of 45" to the plane containing the pivot points of two nozzles, which plane is coincident with, or parallel to, the longitudinal axis of the missile. With this arrangement perfectly complementary movement is achieved between the nozzles and the associated fins in terms of steering control whereby the latter is maximised.

It happens that if two aerodynamic or hydrodynamic surfaces associated with a nozzle are mounted with the pivot axis of one in said plane and the pivot axis of the other at right angles thereto, then only one surface will produce a roll effect, the other being ineffective in this respect so that the steering control is less than maximum.

However, at least the ineffective surface does not positively counteract the roll effect produced by the nozzle and the other surface which can happen with some arrangements of linked thrust vector apparatus and aerodynamic or hydrodynamic surfaces other than in accordance with the present invention.

A satisfactory arrangement is one having two thrust vector control nozzles with two fins or other aerodynamic or hydrodynamic surfaces associated with each nozzle, the pivot axis being at said angle of 45" and this equiangularly spaced round the vehicle. The angular movement of the nozzles about their pivot points may be equal to or different from the consequential movement of the associated fins or other surfaces about their pivot axes.

The linkage means are preferably in the form of simple links extending from the nozzles to the associated aerodynamic or hydrodynamic surfaces.

Each link may be attached at respective ends to a nozzle and one end of crank arm provided on the associated aerodynamic or hydrodynamic surface, the attachments being by way of universal joints.

Guided missiles in accordance with the present invention will now be described in greater detail, by way of example, with reference to the accompanying drawings, in which: Figure 1 is a schematic perspective view of the tail of one missile, and Figure 2 is a schematic view from the rear of the missile of Figure 1, Figure 3 is a side view, with parts broken away, of the tail of another missile, Figures 4 and 5 are views in the direction of the respective arrows IV and V of Figure 3, and Figure 6 is a view in the direction of arrow VI of Figure 5.

Figures 1 and 2 illustrate in a simple, diagrammatic manner one guided missile in accordance with the present invention as applied to a guided missile. The missile comprises a generally cylindrical body 1 of which only the tail is shown in the drawings. A motor (not shown) for the missile is mounted in the body 1 and the exhaust therefrom is ducted by two blast pipes 2 and 3 to respective exhaust nozzles 4 and 5 forming part of thrust vector control apparatus for the missile. The nozzles 4 and 5 are mounted for universal movement about the pivot points 6 in a conventional manner not illustrated, the pivot points 6 lying on a diameter of the missile body as seen in Figure 2. Nozzle movement about the pivot points 6 is effected by four actuators 7 of which only one is shown in Figure 1.

The actuators 7 may be hydraulic or otherwise and comprise single-acting or double-acting cylinders.

The piston 8 of each cylinder is connected to one end of a crank arm 9 the other of which is attached to a shaft 11 on which an aerodynamic fin 12 is mounted. The four fins 12 are mounted on the body 1 and equiangularly spaced around the longitudinal axis 13 of the missile. Each nozzle 4, 5 is linked to two adjacent fins 12 by simple link arms 14 each connected at one end, via a universal joint 15, to the nozzle and at the other end, also via a universal joint 16, to the end of the associated crank arm 9 to which the actuator 7 is attached.

It will be seen that movement of the nozzle 4 and/or 5 provides the necessary guidance of the missile, when the motor is operating, via the thrust vector control apparatus in a conventional manner, i.e. by deflecting the jet stream issuing from the exhaust nozzles. However, movement of the nozzles 4 and 5 also results in rotation of the associated fins 12 so that in the thrust vector control mode, the fin movement complements that of the nozzles to aid the guidance of the missile which is thus even more manoeuvrable than when thrust vector control alone is employed. When the missile motor is no longer operative, the fins 12 can still be controlled as to position by operation of the actuators 7 to give missile guidance in the glide phase, provided the missile has sufficient forward velocity.

In order to simplify the geometry and to make the movement of all four fins 12 complementary to that of the associated nozzles 4 and 5 in terms of missile guidance, the axis of each fin shaft 11 is oriented at 45" to the plane 17 containing the pivot points 6 of the nozzles, which plane passes through the longitudinal axes 13 of the missile, as seen in Figure 2. It is possible to mount one pair of opposed fins 12 in the plane 17 but it is found that in this case one fin of each pair contributes nothing to roll movement of the missile which can be tolerated in some circumstances but it is not normally desirable.

If required, each crank arm 9 can be extended so that the fin is attached intermediate the ends thereof and a second link arm 14' provided to connect the end of the extension to the associated nozzle as indicated in broken lines in respect of one nozzle and fin in Figure 2.

Turning now to Figures 3 to 6, these show a preferred arrangement of a guided missile which again has a generally cylindrical body 20 of which only the tail is shown in the drawings. A motor 21 for the missile is mounted in the body 20 and the exhaust thereof is ducted by two blast pipes 22 and 23 to respective exhaust nozzles 24 and 25 forming part of thrust vector control apparatus for the missile. The nozzles 24 and 25 are mounted for universal movement on part-spherical ends of the respective blast pipes in a conventional manner about pivot points 26 which lie on a vertical diameter of the body 20 as seen in Figure 5. In order to prevent rotary movement of the nozzles 24 and 25 about axes parallel to the longitudinal axis 27 of the missile, two anti-rotation devices are provided of which only one is shown in Figure 4.Each device comprises a pair of ears 30 attached to the associated nozzle 4 or 5, and a pivot pin 28 journal led in bracket 29 fixed to the interior of the missile body 20. The pivot pin 28 has a parallelsided head 28' which is located between the ears 30 and is capable of sliding relative thereto. The pivot pin 28 and the head 28' between them accommodate all movement of the nozzle 4 or 5 relative thereto in all directions except rotation about the longitudinal axis 27 which is prevented by the engagement of the head between the ears 30.

Four adjustable aerodynamic tail fins 31 are mounted on the body 20 at equi-spaced intervals around the longitudinal axis 27, i.e. at 90 to each other, about shafts 32 attached at one end to the respective fins and at the other end to the ends of respective crank arms 33. The fins 31 are connected to the nozzles 24 and 25 by linkage means in a manner such that movement of one nozzle results in movement of two adjacent fins. The linkage means comprise links or rods 34 extending from the other ends of the crank arms 33 and the associated nozzle 24, 25, the links being connected to these components by universal joints.

The nozzles 24, 25 and the fins 31 are moved in unison about the pivot axes 26 and the shafts 32, respectively, by actuator means comprising a hydraulic power supply mounted in the body 20 but not shown in the drawings, four servo valves 35 connected to the power supply and four pairs of single-acting hydraulic cylinders 36, 37 controlled by the respective servo valves. As best seen in Figure 6, the cylinders 36, 37 of each pair are displaced relative to each other longitudinally of the missile body 20 and the rearmost cylinder 36 has its piston 38 connected directly to the associated nozzle by a universal joint. The other cylinder 37 of the pair has its piston 39 connected to the crank arm 33 of one of the fins 31 associated with the nozzle to which the cylinder 36 of that pair is connected.Thus the cylinders 37 are connected to the nozzles by the links 34 through the intermediary of the crank arms 33. A given pair of cylinders 36, 37 is effectively connected (directly or indirectly as explained) to diametrically opposed points in the associated nozzle and each nozzle has two pairs of cylinders connected thereto with the diameter through the connection points of one pair being at right angles to the diameter through the connection points of the other pair. It will be appreciated that in Figure 6 only one pair of cylinders 36, 37 is seen, together with a cylinder 36' of a second pair and a cylinder 37" of a third pair. The cylinders 36, 37 are hydraulically connected to the respective servo valves 35 by passageways drilled in mounting blocks 41 as indicated at 42 in Figure 3.

The pistons 38 and links 34 are connected to the respective nozzles 24, 25 at points 43 which lie in a common plane at right angles to the longitudinal axis 27, which plane is offset by a small amount from the plane containing the nozzle pivot points 26. It is preferable for these two planes to coincide in order to simplify the geometry but in this embodiment space restraints within the missile body has prevented this. It is for the same reason that the cylinders 36, 37 of each pair are displaced longitudinally as it would normally be arranged that all the cylinders are connected directly to the nozzles 24, 25.The moment arm M (Figure 6) of each piston 38 and link 34 relative to the nozzle 24, 25 to which it is connected is greater than the effective length of each crank arm 33 so that for a given angular movement of a nozzle 24, 25, the associated fins 31 move through a greater angle about the fin shafts 32, the ratio of movement of nozzle to fin being approximately 2:3 in this embodiment.

In operation, the missile is launched by firing the motor 21, whereupon hot pressurised gas travels down the blast pipes 22 and 23 and issues from the exhaust nozzles 24 and 25 and propels the missile. With the nozzles 24 and 25 and the fins 31 in the positions shown in full lines throughout Figures 3 to 6, the missile will travel in a straight ahead direction. During this so-called boost phase of the missile flight, the motor operates at maximum thrust, and if the missile has to be manoeuvred either because it is to be engaged in a dog-fight or because it has to be directed to a long distance target, the nozzles 24 and 25 are pivoted in the appropriate direction to give thrust vector control which makes the missile highly manoeuvrable.

Nozzle movement is effected by applying appropriate electrical signals to the servo valves 35 so as to control the supply of hydraulic fluid to the cylinders 36 or 37 which are to be extended and to allow fluid to escape from the cylinder of a pair which is not to be extended but retracted. It will be appreciated that the operation of one pair of cylinders 36, 37 of a given nozzle will result in movement of that nozzle about pivot point 26 in one direction and that operation of the other pair of associated cylinders results in nozzle movement in a direction orthogonal to said one direction, but that operation of both pairs of cylinders provides movement of the nozzle in any desired direction about the pivot point 26. Thus, the universal movement of the nozzles 24 and 25 allows the missile to be guided in pitch, roll and yaw, or any combination thereof, as required.

Since the nozzles 24 and 25 are each linked to two fins 31, then movement of the former about the pivot points 26 results in movement of the fins about their shafts 32 and the arrangement is such that perfectly complementary movement of the nozzles and fins takes place as regards steering or guidance of the missile, whereby the latter is maximised. Figures 3 and 5 indicate in broken lines the movement of the nozzle 24 and an associated Figure 31 from the straight ahead position to another position.

Following the boost phase of the missile flight comes either a coast phase or a glide phase. In the coast phase, the motor 21 produces a reduced thrust compared with the boost phase and steering can still be effected by a combination of thrust vector control through the nozzles 24 and 25 and aerodynamic control through the fins 31 as in the boost phase. In the glide phase, however, the motor 21 is no longer operative so that thrust vector control cannot be effected but provided the forward velocity of the missile is sufficient, aerodynamic control or steering can be implemented by operating the fins 31 through the servo valves 35 as before. The fact that the nozzles 24 and 25 are also moved in this phase of the missile flight is immaterial.

There is thus provided a highly versatile missile having extremely good manoeuvrability in the boost phase which is ideally suited to dog-fight situations or when the missile has to be manoeuvred quickly after launching for other reasons. If the missile is required to engage a long distance target, then specific flight control can be effected over a majority of the flight time by first using the thrust vector control plus aerodynamic control capability followed by aerodynamic control alone.

The invention not only gives rise to such a versatile missile or any other vehicle to which it is applied, but also results in a significant saving in components which would otherwise have to be employed to produce the same versatility. This is because if the nozzles 24 and 25 and the fins 31 were not linked in accordance with the invention then separate actuators for the nozzles and fins would have to be provided. Thus, there is a saving of four actuators one for each fin 31, in the embodiment of Figures 3 to 6 and this means that the missile can be made smaller and more lightweight both of which are very important considerations not only in missile applications but in others.Furthermore, the nozzles and fins are linked in such way that they either give entirely compatible steering movement to the missile vehicle or at least fin movement does not actually counteract the nozzle movement in terms of vehicle guidance. With the combined use of thrust vector control and aerodynamic or hydrodynamic control, improved steering performance is obtained in that a given manoeuvre can be achieved with a smaller angular movement of exhaust nozzle and hence the linked aerodynamic and hydrodynamic surfaces.

It will be clear that changes can be made to the embodiment of Figures 3 to 6 of the drawings without departing from the scope of the invention.

It has already been mentioned that all the cylinders 36 and 37 can operate directly on the nozzles 24 and 25 but alternatively all cylinders could operate on the crank arms 33 or any other intermediary in the linkage between fin and nozzle. Also, the simple link 34 could be replaced by a compound linkage should the circumstances so require. The hydraulic cylinders 36, 37 may be replaced by other forms of actuators such as rotary vane, pneumatic or eiectric actuators, for example.

Instead of, or in addition to, tail fins, canards or other aerodynamic surfaces may be provided which can be linked to the thrust vector control nozzles.

Claims (10)

1. A vehicle fitted with thrust vector control apparatus having a plurality of exhaust nozzles mounted for universal movement about respective pivot points, aerodynamic or hydrodynamic surfaces mounted for pivotal movement about respective pivot axes, and a steering control system comprising actuator means operable to effect pivotal movement of the nozzles, and linkage means extending between the nozzles and said surfaces, the arrangement being such that each nozzle is coupled to at least two of said surfaces and pivotal movement of the nozzle results in movement of the associated surfaces about their respective pivot axes.

2. A vehicle according to claim 1, wherein the pivot axis of each aerodynamic or hydrodynamic surface subtends an angle of 45O to the plane containing the pivot points of two nozzles, which plane is coincident with or parallel to the longitudinal axis of the missile.

3. A vehicle according to claim 1 or 2, wherein the or each linkage is attached at one end to a nozzle and at the other end to a crank arm mounted on a shaft defining the pivot axis of an aerodynamic or hydrodynamic surface.

4. A vehicle according to claim 3, wherein the effective length of each crank arm is less than the moment arm on the nozzle arising out of the point of attachment of the linkage to the associated nozzle relative to the pivot point of the nozzle.

5. A vehicle according to any of the preceding claims, wherein each nozzle has two pairs of actuators associated therewith, one pair arranged to effect nozzle movement in one direction and the other pair arranged to effect nozzle movement in a direction orthogonal to said one direction.

6. A vehicle according to claim 5 when appended to claim 3 or 4, wherein one actuator of a pair is connected directly to the associated nozzle and the other actuator of the pair is connected to the crank arm of one of the fins associated with that nozzle.

7. A vehicle according to claim 6 wherein each pair of actuators is controlled by a servo valve.

8. A vehicle according to any of the preceding claims, wherein the actuator means comprise a hydraulic power supply and hydraulic cylinders hydraulically connected to the power supply.

9. A vehicle according to claim- 8, wherein each cylinder is single acting.

10. A guided missile substantialiy as herein particularly described with reference to Figures 1 and 2 or Figures 3 to 6 of the accompanying drawings.