A method and apparatus for selectively providing electrical power for use in a missile

Abstract: The invention provides an apparatus locatable in a missile tip 56, comprising a turbine impeller 10 operatively connected to an electricity generator 12, and a valve assembly 14 locatable upstream of the impeller 10. In use an induced flow of air passes through a passage 60 defined by the missile tip 56 to drive the turbine 10 and the generator 12. The valve assembly 14 is designed to throttle the flow of air at the maximum flight speeds reached by the missile tip 56. The generator 12 becomes operative again at lower flight speeds to power electronic equipment located in the missile tip 56.

Description: Field of the Invention This invention relates to a method and an apparatus for selectively providing electrical power for use in a missile such as a rocket or an artillery shell. Background to the Invention It has become essential for most modern missile systems to have onboard electrical and electronic systems. Typical of such systems are proximity fuses, guidance and television systems. All these systems obviously require a supply of electrical power. A particular requirement of electrical power supplies for these applications is that they must remain entirely inactive before launch to ensure safe handling by the launching crew. Furthermore the electrical power supply systems must remain reliable even after extended periods of storage. Conventional chemical cells or batteries fail to meet at least one of these requirements, and are generally considered unsuitable for use in missiles. A known method of supplying electrical power to missiles involves providing an air-driven turbine acting on a dynamo located inside the missile. An induced airflow is urged through the turbine during flight. Since no comparable airflow occurs during storage and handling, this type of power supply system makes the missiles intrinsically safe during handling before they are launched. Furthermore the required moving parts are far less susceptible to deterioration during handling and storage then would be the case for a comparable chemical power supply system. A disadvantage of this type of power supply system is encountered, however, when these missiles are fired over long ranges, requiring higher than usual launching speeds. The relative air speeds to which their turbines become exposed along extended flight paths tend to lead to overheating as a result of frictional losses as the missiles travel through the air, while the turbines simultaneously reach excessively high rotational speeds. The present invention seeks to addres this shortcoming. United States patent 2,701, 526 (preamble of claims 1 and 5) describes an air flow regulator for maintaining substantially constant, a quantity of air flowing in a duct and in particular, a regulator for controlling a turbine of an air driven generator. United States patent 3,382,805 describes an air flow actuated delay arming device having a resiliently urged air scoop extendible into a slip-stream to a magnitude determined by the equilibrium of the slip-stream and resilient forces acting thereon. United Kingdom patent 2,007,338 describes a nose portion of a missile for regulating a flow of ram air to a turbine. PCT application 91/11677 and European patent application 0,095,270 describe devices for a missile for limiting the rotational speed of a turbine by regulating the passage of air to the turbine. Summary of the Invention The present invention provides a method of selectively providing electrical power to a missile having a turbine for driving an electricity generator by directing an induced airflow through a passage within the missile towards the turbine while the missile is in flight, characterised by the step of throttling the induced airflow to the turbine by progressively urging a tapered, displaceable spindle axially towards an orifice located in the passage when the speed of the missile exceeds a predetermined limit. The induced airflow may be throttled by progressively urging the tapered, displaceable spindle towards the orifice located in the passage as the rate of the induced airflow increases. The spindle is conveniently displaced by exposing an actuator such as a perforated piston, which is connected to the tapered spindle, to the induced airflow. In a preferred embodiment of the invention the piston is located in the path of the induced airflow, upstream of the spindle and orifice. The invention extends to an apparatus for selectively generating electrical power in a missile, which apparatus includes a turbine impeller locatable in a chamber defined by the missile, the impeller being operatively connected to a rotatable electricity generating means, characterised by the provision of a valve seat defining an orifice, the valve seat being locatable in a passage within the missile for directing an induced airflow through the orifice towards the chamber provided for the turbine when the missile is in flight; a displaceable body having a complementary portion matching the size of the orifice in the valve seat; and resilient biasing means arranged between the valve seat and the displaceable body, the biasing means being so dimensioned as to permit the body to approach the seat at progressively increasing flight speeds by the missile, thereby throttling the airflow to the turbine. The valve seat preferably defines a circular orifice, while the displaceable body conveniently includes an elongate spindle having a frusto-conical tip having a maximum outside diameter exceeding the diameter of the orifice. The displaceable body may further include a piston connected to the spindle, the piston including a perforated plate located upstream of the frusto-conical tip of the spindle, thereby allowing a flow of air through the perforations to urge the tip towards the orifice in the valve seat. The resilient biasing means is conveniently interposed between the perforated plate and the valve seat to ensure that the valve tends to shut when the missile flight speed exceeds a predetermined upper limit. The invention further provides a missile which includes an apparatus of the invention as set out above, comprising; an elongate body tapering in a forward direction towards a tip, the body defining a passage extending rearwardly into the body; a turbine impeller located in a chamber within the body, the chamber being in fluid communication with the passage, and the impeller being so arranged that an induced airflow directed through the passage imparts a rotating motion on the impeller when the missile is in flight; rotatable electricity generating means, operatively connected to the turbine impeller, characterised by the provision of a valve located upstream of the turbine impeller, which valve includes a displaceable body having a portion matching an orifice in a seat located in the passage, and resilient biasing means arranged between the seat and the body, which is so dimensioned that the induced airflow past the body tends to urge it towards the orifice against the bias of the biasing means, shutting the orifice at increasing missile flight speeds. Brief Description of the Drawings The invention is described below by way of example with reference to the accompanying diagrammatic drawings in which Figure 1 shows an exploded 3-dimensional view of a valve assembly forming part of an apparatus according to the invention, portions of the components of the valve assembly being cut away for ease of illustration; and Figure 2 shows an axial section of a missile incorporating the valve assembly of figure 1, illustrating the operation of the apparatus. Description of a Preferred Embodiment of the Invention Figures 1 and 2 show an apparatus according to the invention comprising a turbine impeller 10, an electricity generator 12 and a valve assembly, generally identified by reference numeral 14. The impeller 10 and generator 12 are linked to each other by means of a shaft in known fashion. In use, an axial airflow passage through the impeller 10 causes the shaft to rotate whilst the generator delivers an electrical current through a pair of conductors 16. The valve assembly 14 includes a body 18 of circular cross section defining a first bore 20 and a second bore 22 having a lesser diameter, separated from each other by a stepped shoulder formation 24. The body 18 defines a first annular groove 26 extending radially from the first bore 22 for accommodating a mechanical seal 28. A second similar groove 30 in the second bore 22 accommodates a further mechanical seal 32 (figure 2) of smaller diameter. The valve assembly 14 further includes a piston 34 of circular cross section fitting with slight lateral clearance into the first bore 20. The piston 34 defines a number of passages 36 arranged at regular intervals from each other at a constant distance from the central axis of the piston 34. A spindle 38 is secured to the piston 34 and extends in co-axial relationship to it. The free end of the spindle 38 defines a frusto-conical tip 40 which reduces in cross section towards its free end. A valve seat 42 of circular cross section fits with slight lateral clearance into the first bore 20. The seat 42 defines a central orifice 44 into which the frustoconical tip 40 of the spindle 38 is receivable. The portion of the valve seat 42 defining the orifice 44 is recessed to accommodate a helical spring 46 extending in co-axial relationship to the body 18, piston 34 and seat 42 in their assembled state. The piston 34 is so dimensioned that the spring 46 is receivable into an annular space extending between the spindle 38 and the inner sidewalls of the piston 34. The valve assembly 14 is assembled by successively sliding the piston 34 and the spring 46 past the O-ring 28 located in the groove 26. The valve seat 42 is then similarly inserted into the first bore 20 partially extending past the O-ring 28 and bearing against the spring 46. A spring wire clip 48 is released into an annular groove 50 extending radially from a region of the first bore 20 situated between the first groove 26 and the open end of the valve body 18. The clip 48 is so dimensioned that it partially protrudes inwardly into the first bore 20 once it has been positioned in its associated groove 50. The valve seat 42 defines an inwardly stepped annular shoulder 52 which bears against the protruding portions of the clip 48. The valve assembly 14 is installed stepwise by first locating the valve body 18 in the tip of a missile 54, such as an artillery or rocket shell, having a tapered tip 56 into which a tubular element 58 is receivable, thereby defining a central air passage 60 extending rearwardly from the forward end of the tip. The tubular element 58 terminates in a spool-piece 62 having a bore at one end to which the tubular element fits with marginal clearance, and a nozzle portion at the other end which fits into the second bore 22 of the valve assembly 14. The mechanical seal 32 provides an airtight seal between the valve body 18 and the spool-piece 62. Before completion of the valve assembly 14 a settable foam is injected into the cavity defined between the inner sidewalls of the body of the tapered tip 56, and the outer sidewalls of the tubular element 58, the spool-piece 62 and the valve body 18. This cavity accommodates the electronic components (not shown) which regulate the detonation of the missile associated with the tip 56. The settable foam serves to stabilise the electronic components against shock and possible malfunctioning arising from any handling of the missile. The remaining components of the valve assembly 14 are conveniently installed in the manner described above. This method of installation has the advantage that the risk of any residual foam material finding its way into the moving parts of the valve assembly 14 is substantially diminished. The impeller 10 is located in a chamber 64 within the tip 56, downstream of the completed valve assembly 14. An air exhaust passage 66 extends radially from the chamber 64 towards the outer periphery of the tip 56. The generator 12 is secured to the tip 56 by means of a locking pin 68 fitting into a corresponding recess within the body of the tip. The conductors 16 of the generator 12 are routed within the tip to the electronic components of the missile 54. In use, the missile is launched from a tube or artillery piece (not shown) in known fashion. An induced flow of air enters into the air passage 60 from the leading end of the tip 56. On reaching the valve assembly 14 the airflow is split into separate streams passing through the passages 36 in the piston 34 thereby imparting a frictional drag force on the piston, which tends to urge it rearwardly against the bias of the spring 46. The magnitude of this drag force is determined by the flight speed of the missile 54. Immediately upon launch of the missile the turbine impeller 10 is exposed to a surge of air entering the passage 60 before the spindle 38 forces its frusto-conical tip 40 into the orifice 44 of the valve seat 42, thereby effectively shutting down the impeller 10 and generator 12. The pressure in the stagnant air column in the passage 60 exceeds the pressure within the chamber 64, which is in fluid communication with the outer regions of the missile tip 56. The apparent airspeed past the missile tip 56 generally decreases as it approaches the highest point of its trajectory path, and the static air pressure within the passage 60 decreases correspondingly. The spring 46 hence gradually urges the spindle 38 away from the valve seat 42, thereby admitting a progressively increasing induced airflow through the orifice 44 towards the chamber 64. This in turn causes the turbine impeller 10 and generator 12 to rotate and produce an electrical current for powering the electronic components inside the missile tip 56. The generator 12 sustains this current for a sufficiently long time to allow detonation of the missile at a predetermined height before striking the ground. The action of the valve assembly 14 ensures that the turbine 10 and generator 12 do not reach potentially damaging running speeds, during the flight of the missile tip 56. The valve assembly 14 accordingly enhances the reliability of the turbine impeller 10 and generator 12 by confining their operation to relatively low running speeds for the duration of a time interval during which their operation is crucial. The passages 36 in the piston 34 must be so dimensioned that air flowing through these as the speeds encountered during the launch of the missile 54 imparts a desired drag force on the piston 34. This drag force overcomes the countering force of the helical spring 46, thereby allowing the valve assembly 14 to shut off the airflow to the turbine impeller 10 at what is considered to be the maximum flight speed at which the turbine should be allowed to operate.

Claims: A method of selectively providing electrical power to a missile (54) having a turbine (10) for driving an electricity generator (12) by directing an induced airflow through a passage (60) within the missile (54) towards the turbine (10) while the missile (54) is in flight, characterised by the step of throttling the induced airflow to the turbine (10) by progressively urging a tapered, displaceable spindle (38) axially towards an orifice (44) located in the passage (60) when the speed of the missile (54) exceeds a predetermined limit. A method according to claim 1 in which the induced airflow is throttled by progressively urging the tapered, displaceable spindle (38) towards the orifice (44) located in the passage (60) as the rate of the induced airflow increases. A method according to claim 2 in which the spindle (38) is displaced by exposing a perforated actuator (34), which is connected to the tapered spindle (38), to the induced airflow. A method according to claim 3 in which the actuator (34) is located in the path of the induced airflow, upstream of the spindle (38) and orifice (44). An apparatus for selectively generating electrical power in a missile (54), which apparatus includes a turbine impeller (10) locatable in a chamber (64) defined by the missile (54), the impeller (10) being operatively connected to a rotatable electricity generating means (12), characterised by the provision of a valve seat (42) defining an orifice (44), the valve seat (42) being locatable in a passage (60) within the missile (54) for directing an induced airflow through the orifice (44) towards the chamber (64) provided for the turbine (10) when the missile (54) is in flight; a displaceable body (38) having a complementary portion (40) matching the size of the orifice (44) in the valve seat (42); and resilient biasing means (46) arranged between the valve seat (42) and the displaceable body (38), the biasing means (46) being so dimensioned as to permit the body (38) to approach the seat (42) at progressively increasing flight speeds by the missile (54), thereby throttling the airflow to the turbine (10). An apparatus according to claim 5 in which the valve seat (42) defines a circular orifice (44), while the displaceable body (38) includes an elongate spindle (38) having a frusto-conical tip (40) having a maximum outside diameter exceeding the diameter of the orifice (44). An apparatus according to claim 5 or claim 6 in which the displaceable body (38) includes a piston (34) connected to the spindle (38), the spindle (34) including a perforated plate located upstream of the frusto-conical tip (40) of the spindle (38), thereby in use allowing a flow of air through the perforations (36) to urge the tip (40) towards the orifice (44) in the valve seat (42). An apparatus according to claim 7 in which the resilient biasing means (46) is interposed between the perforated plate (34) and the valve seat (42) to ensure that the valve tends to shut when the missile (54) flight speed exceeds a predetermined upper limit. A missile (54) which includes an apparatus as claimed in any of claim 5 to 8, comprising an elongate body tapering in a forward direction towards a tip (56), the body defining a passage (60) extending rearwardly into the body; a turbine impeller (10) located in a chamber (64) within the body, the chamber (64) being in fluid communication with the passage (60), and the impeller (10) being so arranged that an induced airflow directed through the passage (60) imparts a rotating motion on the impeller (10) when the missile (54) is in flight; rotatable electricity generating means (12), operatively connected to the turbine impeller (10), characterised by the provision of a valve (14) located upstream of the turbine impeller (10), which valve (14) includes a displaceable body (38) having a portion (40) matching an orifice (44) in a seat (42) located in the passage (60), and resilient biasing means (46) arranged between the seat (42) and the body (38), which is so dimensioned that the induced airflow past the body (38) tends to urge it towards the orifice (44) against the bias of the biasing means, shutting the orifice (44) at increasing missile flight speeds.

Inventors: BUDRICKS JACOBUS STEFANUS, BUDRICKS, JACOBUS STEFANUS
Applicants: FUCHS ELECTRONICS PTY LTD, FUCHS ELECTRONICS (PTY) LIMITED
Application year: 1993
Main classification: F42C15
IPC codes: F42C15