THRUSTER PACK FOR MISSILE CONTROLThis invention relates to thruster mechanisms for controlling the attitude of a flying body, such as a guided missile for example.
The invention has particular application to the so-called "Soft Vertical Launch" guided weapon system.
Using this "soft launch" technique, a guided weapon is propelled vertically from a launcher or a container at a relatively low velocity (typically between 30 and 60 msec1) and subsequently manoeuvred into the correct attitude before the main propulsion motor is fired. This system differs from the more conventional vertical launch technique (for example, that employed in our known Seawolf (RTM) system) because the soft-launched missile has insufficient forward velocity to be aerodynamic during said manoeuvre.
However, a soft vertical launch system advantageously removes the requirement for a trainable launcher whilst maintaining the capability for the missile to engage targets incoming from all directions.
It is known to employ thrusters to control missile attitude (i.e. pitch and yaw). Usually a plurality of thrusters are disposed around and integrated with the missile body. See for example, GB-A-2287489. A disadvantage with the known arrangement is that integration of a thruster system can place constraints on the overall missile configuration.
Further, conventional thruster systems require an additional missile roll control mechanism in order to align the missile to the correct roll position before the pitch and yaw control thrusters can be fired.
The present invention, advantageously and in contrast to known thruster mechanisms does not require the missile to be aligned in roll before its deployment.
Accordingly, this invention comprises a thruster mechanism for controlling the attitude of a flying body which incorporates a navigation unit, the thruster mechanism being mountable on and rotatable about the flying body, and including;(i) a rotor portion incorporating at least onethruster unit,(ii) angular rate and position measuring means formonitoring the angular rate and the angularposition of the thruster mechanism with respectto the flying body, and(iii) means for generating a firing signal to initiatefiring of a thruster unit in response to signals received from the navigation unit and the angularrate and position measuring means.
Preferably the thruster mechanism is attached to the rear of the flying body (or missile) in order to maximise the turning moment.
In the preferred embodiment, the mechanism includes a drum-shaped rotor having a plurality of thruster units distributed around its peripheral regions.
At least one thruster unit may be configured to impart a spin torque to the rotor when fired.
For example one or more thruster units may be configured so that the direction of thrust imparted is offset in relation to the axis of rotation of the rotor, thereby imparting a spin torque to the rotor.
Alternatively, or additionally, two spin cartridges may be placed at either end of a diameter of the rotor and configured so that the direction of thrust imparted is offset in relation to the axis of rotation of the rotor thereby imparting a spin torque to the rotor.
Alternative means for spinning up the rotor with respect to the missile may comprise a spring mechanism.
The rotor may be attached to a stator via bearings.
The stator may be fixed over the exhaust nozzle of the missile, the entire mechanism being jettisonable on firing of the main missile motor.
The angular speed and position of the rotor may be monitored by an optical or mechanical shaft encoder, for example.
Following launch of the missile, the thruster units are fired sequentially to pitch the missile in the required direction as each unit is carried into alignment by the spinning rotor.
Because, by virtue of the spinning rotor, the thruster units can be repositioned around the missiles axis whilst in flight, the invention obviates the requirement for the missile to be controlled in the spin (or roll) axis in order to direct the lateral thrust in any given direction.
The rotor may be manufactured from glass reinforced plastics or other structural polymer suitable for economic volume manufacture using a casting or injection moulding technique, for example.
One of the advantages of the invention is that it requires only relatively low power and short duration cartridges to achieve the desired pitching manoeuvre. In comparison with a continuously burning thruster pack, the total explosive content of the rotor is greatly reduced.
A thruster mechanism being of modular construction, and attached to the base of the missile may be subsequently jettisoned when no longer required. Thus the mechanism only needs to withstand a relatively benign environment compared with the missile itself. As the mechanism is a separate item, its use avoids compromising missile integration.
By making the thruster mechanism discardable after use, it does not make any contribution to the mass of the missile in flight, which is a further advantage. Thus in contrast to a conventional integrated thruster mechanism, it enhances potential performance of the missile in terms of range and terminal velocity.
A drogue arresting device may be included in the thruster mechanism for controlling its descent after being jettisoned from the missile.
An embodiment of the invention will now be described with reference to the drawings of which:Figure 1 is a side view of a missile incorporating a thruster mechanism in accordance with the invention;Figure 2 is a side view of the thruster mechanism ofFigure 1;Figure 3 is a cross-sectional view along a line III-III of Figure 2; Figure 4 is a schematic block diagram of the thruster mechanism control circuitry; andFigure 5 is a series of diagrams illustrating the launch sequence of the missile of Figure 1.
In Figure 1 a thruster pack module 1 is shown attached to the rear portion of a missile 2. The missile 2 is positioned within a launch tube 3.
The thruster pack module 1 is preferentially located at the rear of the missile in order that the turning motion it provides to the missile, for a given thrust, is maximised.
The calibre of the thruster pack is chosen to be large enough to provide an obturating base for the launch tube 3.
Referring now to Figures 2 and 3, the mechanical parts of the thruster pack module comprise a stator 4 and a rotor 5. The stator 4, which is essentially an annular ring, is fitted over the nozzle 6 of the missile's main propulsion motor. The rotor 5 is essentially a cylindrical drum with a central hole in which the stator 4 sits, separated from the rotor and coupled thereto via needle bearings 7. The rotor is manufactured from glass reinforced plastics and contains nine single shot lateral thrust cartridges 8 and two smaller spin cartridges 9. The two spin cartridges 9 are symmetrically placed at either end of a diameter of the rotor, in order to spin the rotor 5 when fired in unison.
The direction of thrust provided by the lateral thrust cartridges 8 is arranged (by appropriate positioning of the cartridges 8) to be slightly offset in relation to the axis of rotation of the rotor 5. In this way, an additional small spin torque is imparted to the rotor 5 when each lateral thrust cartridge 8 is fired.
Each lateral thrust cartridge 8 and each spin cartridge 9 is filled with propellant 10 and incorporates a nozzle 11 and an independent means of ignition (not shown).
The rotor 5 is also provided with a shaft encoder 12 which monitors the angular rate of the rotor 5 and its angular position relative to missile 1.
The electrical output from the shaft encoder 12 is applied to a firing control module 13 which is contained within the rotor and whose operation is to be described herebelow with reference to Figure 4.
A slip ring (not shown) fitted to the rotor 5 interfaces the firing control module 13 with an inertial navigation unit (INU) (see Figure 4) carried by the missile 1.
The entire rotor/stator pack of Figure 3 is adapted to be jettisoned on firing of the missile's main motor.
Turning now to Figure 4 the firing control module 13 receives data inputs from the INU 14 via a digital serial data link. This data comprises demanded and achieved missile attitude. The module 13 also receives angular rate and angular position data from the shaft encoder 12.
Pre-programmed into the module 13 are the thrust characteristics of each of the thrust cartridges 8, (which do not necessarily need to be the same value).
The firing control module 13, together with the shaft encoder 12 and INU 14 embody a semi-closed loop, adaptive digital servo whose function is to output a firing signal on line 15 to the appropriate cartridge 8 at the correct instant in time in order to direct the missile 1 in the desired direction.
In operation, (see Figure 5) the missile 1 is expelled from the launch tube 3 with sufficient velocity to achieve the turnover altitude.
Immediately following launch-tube exit, the two spin cartridges 9 are simultaneously ignited (Figure 5a) in response to a firing signal from the firing control module, in order to spin the rotor 5 about the missile's longitudinal axis.
After spin-up of the rotor 5, the firing control module 13 initiates the firing of several lateral thrust cartridges 8 in succession (Figures 5b, 5c, sod). As there is no need to align the missile in roll, the turnover manoeuvre can be commenced almost immediately after launch tube exit.
Firing of the thrust cartridges 8 in order to commence the turnover manoeuvre has the effect of pitching the missile 1 into the required attitude demanded by the INU 14.
The additional small rolling couple imparted to the rotor 5 by the firing of the thrust cartridges 8 at least maintains if not increases spin rate of the rotor 5. Rotor spin rate and its position relative to the body of the missile 1 is constantly monitored by the shaft encoder 12. Using data from the encoder 12 and data relating to demanded and achieved attitude from the INU 14, the firing control module 13 can progressively compensate for yaw errors and sustain or modify (as necessary) missile pitch rate by firing the remaining thrusters 8.
When acceptable limits for pitch rate, direction, attitude and time-of-flight are achieved, the missile's main motor can be fired. This act blows off the thruster pack 1, jettisoning both rotor 5 and stator 4 clear of the missile (Figure 5e).