WO2018234766A1

WO2018234766A1 - AIRCRAFT PROPELLER

Info

Publication number: WO2018234766A1
Application number: PCT/GB2018/051671
Authority: WO
Inventors: Roger Mark Sloman
Original assignee: Active Vtol Crash Prevention Ltd
Current assignee: Active Vtol Crash Prevention Ltd
Priority date: 2017-06-22
Filing date: 2018-06-15
Publication date: 2018-12-27
Anticipated expiration: 2019-12-22
Also published as: GB201710011D0

Abstract

A rocket motor for a vertical take-off and landing (VTOL) aircraft is provided, along with control circuitry and computer programs for controlling the rocket motor, the methods of using the rocket motor. The rocket motor may define an elongate chamber for storing propellant, the elongate chamber having a length dimension, a width dimension and a depth dimension, wherein the length dimension is greater than the width dimension and greater than the depth dimension. The rocket motor may also comprise propellant and at least one fluid injector. The propellant may be located in the elongate chamber and be configured to enable the rocket motor to generate thrust in a direction that is substantially perpendicular to the length dimension of the elongate chamber by causing ejection of gas through a plurality of gas efflux apertures in the casing. The fluid injector(s) may be for injecting a fluid fire suppressant into the elongate chamber.

Description

TITLE

AIRCRAFT THRUSTER TECHNOLOGICAL FIELD

Embodiments of the present invention relate to rocketry. In particular, they relate to rocket motors for vertical take-off and landing (VTOL) aircraft. BACKGROUND

VTOL aircraft include helicopters, aeroplanes and drones, for example. A benefit of a VTOL aircraft is that it can take-off and land in a relatively confined space in comparison with a conventional take-off and landing aircraft.

BRIEF SUMMARY

According to various, but not necessarily all, embodiments of the invention there is provided a rocket motor for a vertical take-off and landing (VTOL) aircraft, the rocket motor comprising: a casing defining an elongate chamber for storing propellant, the elongate chamber having a length dimension, a width dimension and a depth dimension, wherein the length dimension is greater than the width dimension and greater than the depth dimension; propellant, located in the elongate chamber, configured to enable the rocket motor to generate thrust in a direction that is substantially perpendicular to the length dimension of the elongate chamber by causing ejection of gas through a plurality of gas efflux apertures in the casing; and at least one fluid injector for injecting a fluid fire suppressant into the elongate chamber.

The at least one fluid injector may be for injecting a fluid fire suppressant into the elongate chamber, while thrust is being generated by the rocket motor, in order to increase the thrust. The at least one fluid injector may be for injecting a fluid fire suppressant into the elongate chamber after thrust generation by the rocket motor has ceased. The at least one fluid injector may be for injecting the fluid fire suppressant into a region of the elongate chamber that is located between the propellant and the plurality of gas efflux apertures. The casing may further comprise a conduit, located in the region of the elongate chamber between the propellant and the plurality of gas efflux apertures in the casing, into which the fluid fire suppressant is injected.

The conduit may be an elongate conduit extending along the length dimension of the elongate chamber. The conduit may comprise a plurality of fluid diffusing apertures arranged to release the fluid fire suppressant from the conduit within the region of the elongate chamber located between the propellant and the plurality of gas efflux apertures. The conduit may be a pyrotechnic initiation conduit that enables ignition of the propellant, following ignition of the initiator.

The fluid fire suppressant comprises: one or more inert gases, carbon dioxide, water, anti-freeze, halon, and/or one or more halocarbon-based agents. The at least one fluid injector may comprise a control valve for controlling a flow rate of the fluid fire suppressant into the elongate chamber.

The rocket motor may further comprise: one or more heaters arranged to heat the fluid fire suppressant. The one or more heaters may be configured to heat the fluid fire suppressant prior to injection of the fluid fire suppressant into the elongate chamber.

According to various, but not necessarily all, embodiments of the invention there is provided a VTOL aircraft comprising the rocket motor described above.

According to various, but not necessarily all, embodiments of the invention there is provided an apparatus comprising the rocket motor described above and further comprising: control circuitry for controlling the at least one fluid injector. The control circuitry may be configured to control the at least one fluid injector to inject the fluid fire suppressant into the elongate chamber, while thrust is being generated by the rocket motor. The control circuitry may be configured to continue to control the at least one fluid injector to inject the fluid fire suppressant into the elongate chamber, after thrust generation by the rocket motor has ceased. The apparatus may further comprise sensor circuitry for sensing attitude, altitude, drift and/or descent rate of a VTOL aircraft. The control circuitry may be configured to control the at least one fluid injector based, at least in part, on one or more inputs from the sensor circuitry.

According to various, but not necessarily all, embodiments of the invention there is provided a VTOL aircraft comprising the apparatus described above.

According to various, but not necessarily all, embodiments of the invention there is provided a method, comprising: analysing one or more inputs quantifying an attitude, altitude, drift and/or descent rate of a VTOL aircraft; and controlling a flow rate of fluid fire suppressant into an elongate chamber of a rocket motor based, at least in part, on the analysis of the inputs. According to various, but not necessarily all, embodiments of the invention there is provided computer program code that, when executed by at least one processor, causes the method described above to be performed.

The computer program code may be stored on a non-transitory computer readable medium.

According to various, but not necessarily all, embodiments of the invention there is provided an apparatus, comprising: at least one processor; and memory storing computer program code that, when executed by the at least one processor, causes the apparatus to: analyse one or more inputs quantifying an attitude, altitude, drift and/or descent rate of a VTOL aircraft; and control a flow rate of fluid fire suppressant into an elongate chamber of a rocket motor based, at least in part, on the analysis of the inputs. According to various, but not necessarily all, embodiments of the invention there is provided a rocket motor for a vertical take-off and landing (VTOL) aircraft, the rocket motor comprising: a casing defining an elongate chamber for storing propellant, the elongate chamber having a length dimension, a width dimension and a depth dimension, wherein the length dimension is greater than the width dimension and greater than the depth dimension; propellant, located in the elongate chamber, configured to enable the rocket motor to generate thrust in a direction that is substantially perpendicular to the length dimension of the elongate chamber by causing ejection of gas through a plurality of gas efflux apertures in the casing; an elongate conduit, located in a region of the elongate chamber between the propellant and the plurality of gas efflux apertures in the casing, comprising a plurality of fluid diffusing apertures; and at least one fluid injector for injecting a fluid into the elongate conduit.

The at least one fluid injector may be for injecting a fluid into the elongate conduit, while thrust is being generated by the rocket motor, in order to increase the thrust.

The fluid may be a fire suppressant. The fluid may be a liquid propellant.

According to various, but not necessarily all, embodiments of the invention there is provided a method, comprising: analysing one or more inputs quantifying an attitude, altitude, drift and/or descent rate of a VTOL aircraft; and controlling a flow rate of fluid into an elongate conduit extending along an elongate chamber of a rocket motor based, at least in part, on the analysis of the inputs.

The elongate conduit may be located between propellant in the elongate chamber of the rocket motor and a plurality of gas efflux apertures in a casing of the rocket motor.

According to various, but not necessarily all, embodiments of the invention there is provided computer program code that, when executed by at least one processor, causes the method described above to be performed.

According to various, but not necessarily all, embodiments of the invention there is provided an apparatus, comprising: at least one processor; and memory storing computer program code that, when executed by the at least one processor, causes the apparatus to: analyse one or more inputs quantifying an attitude, altitude, drift and/or descent rate of a VTOL aircraft; and control a flow rate of fluid into an elongate conduit extending along an elongate chamber of a rocket motor based, at least in part, on the analysis of the inputs. According to various, but not necessarily all, embodiments of the invention there is provided a rocket motor for a vertical take-off and landing (VTOL) aircraft, the rocket motor comprising: a casing defining a chamber for storing propellant, the propellant being for generating thrust by causing ejection of gas through a plurality of gas efflux apertures in the casing; and at least one fluid injector for injecting a fluid fire suppressant into the elongate chamber.

According to various, but not necessarily all, embodiments of the invention there is provided examples as claimed in the appended claims.

BRIEF DESCRIPTION

For a better understanding of various examples that are useful for understanding the detailed description, reference will now be made by way of example only to the accompanying drawings in which:

fig. 1 illustrates control circuitry;

fig. 2 illustrates an apparatus comprising the control circuitry, sensor circuitry and one or more rocket motors;

fig. 3A illustrates a first cross-sectional schematic of a first embodiment of a rocket motor;

fig. 3B illustrates a second cross-sectional schematic of the first embodiment of the rocket motor;

fig. 4 illustrates a flow chart of a method;

fig. 5 illustrates a cross-sectional schematic of a second embodiment of the rocket motor;

fig. 6 illustrates a cross-sectional schematic of a third embodiment of the rocket motor; fig. 7A illustrates a first cross-sectional schematic of a fourth embodiment of the rocket motor; and

fig. 7B illustrates a second cross-sectional schematic of a fourth embodiment of the rocket motor.

DETAILED DESCRIPTION Embodiments of the invention relate to the control and use of one or more rocket motors. The rocket motor(s) may, for example, form part of a VTOL aircraft. The rocket motors are configured to provide controllable, variable thrust as required/desired, for example when landing the VTOL aircraft in an emergency situation when normal control has been lost. Alternatively or additionally, the rocket motors are configured to provide a means of reducing or eliminating any fire or burning after thrust generation by the rocket motor has ceased.

In some embodiments, a fluid fire suppressant is injected into a chamber of the rocket motor while the rocket motor is producing thrust and continues to be injected after thrust generation has ceased. Injecting the fluid fire suppressant into the chamber of the rocket motor while the rocket motor is producing thrust causes an increase in the thrust provided by the rocket motor. Injection of the fluid fire suppressant into the chamber after thrust generation has ceased may help to reduce or eliminate any fire or burning (e.g. after the VTOL aircraft has landed).

Fig. 1 illustrates an apparatus 10 in the form of control circuitry. The control circuitry 10 may, for example, be a chip or a chipset. The control circuitry 10 may be implemented in hardware alone, have certain aspects in software including firmware alone or can be a combination of hardware and software (including firmware).

The illustrated control circuitry 10 comprises at least one processor 12 and at least one memory 14. The processor 12 is configured to read from and write to the memory 14. The processor 12 may also comprise an output interface via which data and/or commands are output by the processor 12 and an input interface via which data and/or commands are input to the processor 12.

The memory 14 stores a computer program 16 comprising computer program instructions (computer program code) 18 that controls the operation of the apparatus illustrated in fig. 2, when loaded into the processor 12. The computer program instructions 18, of the computer program 16, provide the logic and routines that enables the apparatus to perform the method illustrated in fig. 4. The processor 12 by reading the memory 14 is able to load and execute the computer program 16. As illustrated in fig. 2, the computer program 16 may arrive at the apparatus 20 via any suitable delivery mechanism 40. The delivery mechanism 40 may be, for example, a non-transitory computer-readable storage medium, a computer program product, a memory device, a record medium such as a compact disc read-only memory (CD- ROM) or digital versatile disc (DVD), an article of manufacture that tangibly embodies the computer program 16. The delivery mechanism may be a signal configured to reliably transfer the computer program 16. The apparatus 20 may propagate or transmit the computer program 16 as a computer data signal. Although the memory 14 is illustrated as a single component/circuitry it may be implemented as one or more separate components/circuitry some or all of which may be integrated/removable and/or may provide permanent/semi-permanent/ dynamic/cached storage. Although the processor 12 is illustrated as a single component/circuitry it may be implemented as one or more separate components/circuitry some or all of which may be integrated/removable. The processor 12 may be a single core or multi-core processor. Fig. 2 illustrates an apparatus 20 comprising one or more rocket motors 100, sensor circuitry 22 and the control circuitry 10 illustrated in fig. 1. A VTOL aircraft may be provided which includes the apparatus 20.

The processor 12 is configured to receive inputs from the sensor circuitry 22 and configured to control the rocket motor(s) 100 based, at least in part, on those inputs.

The sensor circuitry 22 may, for example, comprise one or more attitude sensors, one or more altimeters and/or one or more variometers. The attitude sensor(s) is/are configured to sense the attitude of a VTOL aircraft and may, for example, include one or more gyroscopes, one or more magnetometers, a light detection and ranging (LIDAR) system, and/or one or more image sensors.

The altimeter(s) is/are configured to sense the altitude of a VTOL aircraft and may include one or more barometric altimeters and/or one or more radar altimeters. The variometer(s) is/are configured to sense a rate of descent/rate of change of altitude of a VTOL aircraft.

The rocket motors 100, sensor circuitry 22, processor 12 and memory 14 are operationally coupled and any number or combination of intervening elements can exist between them (including no intervening elements). For example, intervening circuitry may be present between the processor 12 and each rocket motor 100, and between the processor 12 and the sensor circuitry 22. Fig. 3A illustrates a first cross-sectional schematic of a first embodiment 100a of a rocket motor 100. Fig. 3B illustrates a second cross-sectional schematic of the first embodiment 100a of the rocket motor 100, through the line X-X in fig. 3A.

The illustrated rocket motor 100a comprises a casing 102 defining an elongate chamber 103 for storing propellant 104. The propellant 104 may be a solid propellant, and could, for example, be in the form of fins or pellets. An insulating liner 1 12 may be provided in the elongate chamber 103 to insulate the casing 102 from at least some of the heat generated when the propellant 104 inside the elongate chamber 103 is burnt. The rocket motor 100a may be integrated to the structure of a VTOL aircraft. For example, the casing 102 may be integrated with (i.e. form part of) the primary structure of a VTOL aircraft. The primary structure is considered to be the structural portion(s) of the VTOL aircraft which would cause structural collapse and/or inflight loss of control if it/they were to fail.

The elongate chamber 103 is defined by a plurality of walls of the casing and has a length dimension L, a width dimension W and a depth dimension D. The length dimension L is greater than the width dimension W and greater than the depth dimension D. For this reason, the rocket motor may be referred to as a "linear rocket motor".

In some examples, the length dimension L may be at least 1.25 times as great as the width dimension W and/or at least 1.25 times as great as the depth dimension D. In other examples, the length dimension L may be at least twice as great as the width dimension W and/or at least twice as great as the depth dimension D. In some further examples, the length dimension L may be at least five times as great as the width dimension W and/or at least five times as great as the depth dimension D. It might be that the length dimension L is much longer than the width dimension W and/or the depth dimension D, such as 50 times greater or more.

The length dimension L is orthogonal to the width dimension W and the depth dimension D. The width dimension W is orthogonal to the depth dimension D.

The rocket motor 100a further comprises a fluid storage chamber 108, at least one fluid injector 113, a conduit 106 and an initiator/ignitor 1 10.

In the illustrated example, the fluid storage chamber 108 is defined by the casing 102 of the rocket motor 100a and is located adjacent the elongate chamber 103 for storing propellant 104.

The fluid storage chamber 108 is for storing a fluid (i.e. a liquid or a gas) which is for injection into the elongate chamber 103. In some implementations, the fluid is a fluid fire suppressant. The fire suppressant might be a liquid or a gas. For example, the fire suppressant might be or include: one or more inert gases (e.g. nitrogen and/or argon), carbon dioxide, water, anti-freeze, halon, one or more halocarbon-based agents, and their equivalents. In other examples, the fluid is not a fire suppressant and might, for example, be a liquid propellant (e.g. an oxidiser, fuel or monopropellant).

One or more passages enable fluid to be communicated from the fluid storage chamber 108 to the elongate chamber 103, as will be described in further detail below. In other examples, however, the fluid storage chamber 108 may be spaced from the elongate chamber 103 with a conduit therebetween to allow the communication of fluid.

The at least one fluid injector 1 13 is for injecting the fluid stored in the fluid storage chamber 108 into the elongate chamber 103.

The conduit 106 extends through the elongate chamber 103, in the length dimension L, and is elongate in shape. The elongate chamber 103 includes a first region 103a in which propellant 104 is located and a second region 103b that, in the example, is a void. The casing 102 includes a plurality of gas efflux apertures via which gas is ejected from the rocket motor 100a in operation. The conduit 106 is at least partially located between the first region 103a and the second region 103b or, alternatively stated, between the propellant 104 and the gas efflux apertures 109. The conduit 106 may define a void extending along the length dimension L of the elongate chamber 103. The void in the conduit 110 may be present along (at least) a majority of the length of the conduit 106 and (at least) a majority of the length of the elongate chamber 103.

The conduit 106 has two main functions. The conduit 106 enables the propellant 104 in the elongate chamber 103 to be ignited and it enables fluid from the fluid storage chamber 108 to be dispersed along substantially the whole of the length dimension L of the elongate chamber 103. In order to enable the conduit 106 to perform this functionality, pyrotechnic material in addition to that in the initiator 1 10 might be contained within the conduit 106 (prior to ignition of the initiator 1 10) and the conduit 106 comprises a plurality of apertures 107 in its outer surface. The apertures 107 are positioned along the length of the conduit 106. Each of the apertures 107 may be directed towards the (face of the casing 102 comprising the) gas efflux apertures 109.

In use, the control circuitry 10 activates the rocket motor 100a by providing a signal which causes initiation of the rocket motor 100a via an electrical connection 11 1 to the initiator 1 10. The signal provided by the control circuitry 10 causes igniferous material (such as mixed boron-potassium nitrate) of the initiator 1 10 to ignite, which in turn causes hot gas and burning igniferous material (such as boron particles) to pass through the elongate conduit 106. The burning igniferous material is channelled along the conduit 106 and emitted from the apertures 107 in the conduit 106, causing ignition of the propellant 104.

The conduit 106 can be considered to be a pyrotechnic initiation conduit 106 that enables ignition of the propellant 104 because it channels burning igniferous material from the initiator 110 towards the propellant 104, and/or because it comprises igniferous/pyrotechnic material for ignition by the initiator 1 10 and ignition of that igniferous/pyrotechnic material causes or aids ignition of the propellant 104. The igniferous/pyrotechnic material contained in the conduit 106 may, for example, be mixed boron-potassium nitrate, for example). Preferably the initiator 1 10 is of the exploding foil initiator type which has inherent safety characteristics especially applicable to aerospace applications.

In some implementations, one or more covers may be provided which are arranged to cover the gas efflux apertures 109 when the rocket motor 100a is activated, in order to prevent the burning igniferous material from exiting the casing 103 and aid ignition of the propellant 104.

Ignition of the propellant 104 causes a large volume of gas to exit the elongate chamber 103 via the gas efflux apertures 109 in the direction indicated by the arrow labelled with the reference numeral 90 in fig. 3A. The pressure build-up inside the elongate chamber 103 causes the one or more covers, if present, to be removed (e.g. by breaking up or by being blown away). Ejection of the gas via the gas efflux apertures 109 generates a thrust in the direction of the arrow labelled with the reference numeral 85 in fig. 3A. As can be seen in fig. 3A, thrust is generated in a direction that is substantially perpendicular to the length dimension of the elongate chamber 103. The control circuitry 10 is able to cause the rocket motor 100a to generate additional thrust, while thrust is being generated by the rocket motor 100a, by causing at least some of the fluid in the fluid storage chamber 108 to be injected into the elongate chamber 103 of the casing 102. In order to do this, the control circuitry 10 causes a signal to be provided to the one or more fluid injectors 113 via an electrical connection 101 , which in turn causes the fluid injector 113 to open one or more passages between the conduit 106 and the fluid storage chamber 108. The opening of the passage(s) may result in fluid exiting the fluid storage chamber 108 and entering the conduit 106. The fluid injector(s) 1 13 may, for example, be control valve(s). In some examples, the fluid injector 1 14 might be configured to (actively) urge the fluid from the fluid storage chamber 108 into the conduit 106.

For example, the fluid in the fluid storage chamber 108 might be a liquid that would be in a liquid state at ambient temperature and pressure, or a liquefied gas that would be in a gaseous state at ambient temperature and pressure or a compressed gas. In the latter two cases, the fluid injector 1 13 may be configured to merely open a passage to cause the liquefied/compressed gas to exhaust from the fluid storage chamber 108 and into the conduit 106, which would occur due to the pressure differential between the fluid storage chamber 108 and the void in the conduit 106. In the former case, the fluid injector 113 may be configured to (actively) urge the liquid from the fluid storage chamber 108 using an additional compressed gas cylinder (e.g. containing an inert gas and a release mechanism) or other mechanism which drives a piston in the fluid storage chamber 108 to push the liquid into the conduit 106. An embodiment of the rocket motor 100 which contains a piston 99 is described below in relation to fig. 5. The apertures 107 in the conduit 106 act as fluid diffusing apertures which release the fluid from the conduit 106 and into the second region 103b of the elongate chamber 103, located between the propellant 104 and the plurality of gas efflux apertures 109.

The injection of the fluid (from the fluid storage chamber 108) into the second region 103b of the elongate chamber 103 increases the overall mass of fluid that is present in the second region 103b of the elongate chamber 103 and therefore results in a greater mass of fluid being ejected via the gas efflux apertures 109 than was previously the case. This increases the amount of thrust that is provided. In some implementations, the at least one fluid injector 113 is or comprises at least one control valve for controlling a flow rate of the fluid from the fluid storage chamber 108 to the elongate conduit 106. For instance, the control valve(s) may open one or more passages to a greater or lesser extent in order to control the flow rate of the fluid. A greater injection of fluid into the conduit 106 will result in a greater increase in thrust being provided.

The injection of the fluid into the second region 103b of the elongate chamber 103 increase the burning rate of the propellant 104 because the pressure inside the elongate chamber 103 has increased, thereby increasing the amount of thrust that is provided by the rocket motor 100a.

As mentioned above, the fluid that is stored in the fluid storage chamber 108 may be a fluid fire suppressant including one or more of: one or more inert gases (e.g. nitrogen and/or argon), carbon dioxide, water, anti-freeze, halon, one or more halocarbon- based agents, and their equivalents. An injection of a fluid fire suppressant into the conduit 106 might not suppress the burning of the propellant 104 by any significant amount while thrust is being generated by the rocket motor 100a. This is in part due to the amount of propellant 104 that is burning during thrust generation, the high heat of the propellant 104, and the directivity of the apertures 107 in the conduit 106 (i.e. directed away from the propellant 104 and towards the gas efflux apertures 109).

The control circuitry 10 may be configured to control the fluid injector 1 13 to continue to inject the fluid fire suppressant into the conduit 106 (and therefore into the elongate chamber 103) after thrust generation by the rocket motor 100a has ceased. Advantageously, this may help to extinguish any remaining fire/burning within the elongate chamber 103 and possibly help to extinguish any remaining fire/burning outside the elongate chamber 103 as the fluid fire suppressant exits the gas efflux apertures 109. In some implementations, the rocket motor 100a may be used by a VTOL aircraft to make a successful landing in an emergency situation. For example, as the VTOL aircraft descends, the rocket motor 100a may provide a thrust having an upwards component that reduces the rate of descent of the VTOL aircraft and therefore enables it to land safely. In the method illustrated in fig. 4, in block 401 the control circuitry 10 analyses inputs from the sensor circuitry 22. The inputs may, for example, quantify an attitude, altitude, drift, and/or descent rate of the VTOL aircraft. The control circuitry 10 determines, from the analysis, how to control the fluid flow rate from the fluid storage chamber 108 into the conduct 106 (and therefore into the elongate chamber 103) of the rocket motor 100a, and does so in block 402 of fig. 4.

In some examples, the control circuitry 10 may control the fluid flow rate in accordance with a plurality of preset flow rate levels that are selectable by the control circuitry 10 based on inputs from the sensor circuitry 22. In other examples, the control circuitry 10 may control the flow rate in a continuously variable manner.

The control circuitry 10 is able to vary the thrust that is provided by the rocket motor 100a as the VTOL aircraft descends, to provide a safe landing. Furthermore, the continued ejection of a fluid fire suppressant into the elongate chamber 103 of the rocket motor 100a after thrust generation has ceased reduces or eliminates fire risk by extinguishing any fire/burning within the elongate chamber 103 and potentially also fire/burning outside the elongate chamber 103.

Fig. 5 illustrates a cross-sectional schematic of a second embodiment 100b of the rocket motor 100. The second embodiment 100b differs from the first embodiment 100a in that a piston 99 is provided within the fluid storage chamber 108 to urge the fluid from the fluid storage chamber 108 into the elongate conduit 106. In this embodiment, as explained above, an additional compressed gas cylinder (e.g. containing an inert gas and release mechanism) or a different mechanism might be used to drive the piston 99 in the fluid storage chamber 108 to push the fluid into the conduit 106. The second embodiment of the rocket motor 100b may be preferable if, for example, the fluid stored in the fluid storage chamber 108 is a liquid at ambient pressure and temperature rather than a compressed/liquefied gas.

Fig. 6 illustrates a cross-sectional schematic of a third embodiment 100c of the rocket motor 100. The third embodiment 100c may or may not include the piston 99 illustrated in fig. 5, although the piston 99 is not shown in fig. 6.

The third embodiment 100c differs from the first embodiment 100a and the second embodiment 100b in that it comprises one or more optional heaters 114 which are arranged to heat the fluid that has exited the fluid storage chamber 108, prior to the entry of that fluid into the elongate conduit 106 and the elongate chamber 103. By heating the fluid prior to its entry into the elongate chamber 103, it is thought that the efficiency of the rocket motor 100 may be improved. This is because the temperature of the fluid upon entry into the second region 103b of the elongate chamber 103 will be closer to the temperature of the hot gas in the second region 103b and the propellant 104 than would otherwise be the case without the pre-heating.

The one or more heaters 114 system may comprise a pyrotechnic material (e.g. contained within a matrix of a heat exchange system) to provide additional heat to increase the temperature of the fluid before it enters the elongate chamber 103. This pyrotechnic material might be initiated/ignited separately from the main initiator 110, at the same time or a different time than initiation/ignition of the main initiator 110. Fig. 7 A illustrates a first cross-sectional schematic of a fourth embodiment 100d of the rocket motor 100. Fig. 7B illustrates a second cross-sectional schematic of the fourth embodiment 100d, through the line Y-Y in fig. 7A. The fourth embodiment 100d differs from the first, second and third embodiments 100a, 100b, 100c in that the fluid storage chamber 108 is provided by a tank 1 16 that is external to the casing 102 of the rocket motor 100d. A plurality of mechanical connectors 1 15 are provided to connect the tank 1 16 to the body of an aircraft and a further plurality of mechanical connectors 117 are provided to connect the tank 116 to the casing 102 of the rocket motor 100d.

The fourth embodiment 100d also differs from the other illustrated examples in that multiple fluid injectors/control valves 1 13 are provided to inject fluid into the elongate conduit 106, rather than a single fluid injector/control valve 1 13.

In some examples, the tank 116 may be smaller than that illustrated in figs 7A and 7B. It might be that one or more smaller fluid storage chambers 108 are provided in a tank.

References to 'computer program' or a 'control circuitry', 'computer', 'processor' etc. should be understood to encompass not only computers having different architectures such as single/multi-processor architectures and sequential (Von Neumann)/parallel architectures but also specialized circuits such as field-programmable gate arrays (FPGA), application specific circuits (ASIC), signal processing devices and other processing circuitry. References to computer program, instructions, code etc. should be understood to encompass software for a programmable processor or firmware such as, for example, the programmable content of a hardware device whether instructions for a processor, or configuration settings for a fixed-function device, gate array or programmable logic device etc. The blocks illustrated in fig. 4 may represent steps in a method and/or sections of code in the computer program 16. The illustration of a particular order to the blocks does not necessarily imply that there is a required or preferred order for the blocks and the order and arrangement of the block may be varied. Furthermore, it may be possible for some blocks to be omitted. Where a structural feature has been described, it may be replaced by means for performing one or more of the functions of the structural feature whether that function or those functions are explicitly or implicitly described. Although embodiments of the present invention have been described in the preceding paragraphs with reference to various examples, it should be appreciated that modifications to the examples given can be made without departing from the scope of the invention as claimed. For example, the rocket motor 100 may be suitable for use by vehicles or devices other than a VTOL aircraft.

Rocket motors 100 may be positioned either side of the centreline of an aircraft (e.g. one the wings of the aircraft) and/or either side of the centre of gravity of the aircraft if the rocket motors 100 are fore and aft along the centreline). A rocket motor 100 might be rotatable, relative to the attitude of an aircraft. The rotation might be about an axis that is parallel with (and possibly coincident with) the length dimension L of the rocket motor 100.

Rotation of the rocket motor 100 might be controlled by the control circuitry 10. Rotation of the rocket motor 100 enables a groundwards component of the thrust 85 that is generated by the rocket motor 100 to be controlled. For example, the control circuitry 10 might rotate a rocket motor 100 to reduce the groundwards component of the thrust 85 (e.g. from a situation in which the thrust 85 is substantially perpendicular to a situation in which the thrust 85 is not substantially perpendicular to ground). The control circuitry 10 might, for instance, rotate a plurality of rocket motors 100 in this manner while the aircraft is performing an emergency landing, in order to perform an appropriately controlled landing (e.g. such that the horizontal component of the generated thrust on one side of the centreline of the aircraft is counteracted by the horizontal component of the generated thrust on the other side of the centreline of the aircraft, and the overall vertical component of the thrust is reduced compared with prior to the rotation). Alternatively or additionally, rotation of one or more rocket motors 100 in this manner may be performed to control the attitude, roll and/or pitch of an aircraft while the aircraft is airborne. If the rocket motors 100 are used to control the attitude, roll and pitch of an aircraft, the aircraft might include at least four rocket motors 100. Features described in the preceding description may be used in combinations other than the combinations explicitly described.

Although functions have been described with reference to certain features, those functions may be performable by other features whether described or not.

Although features have been described with reference to certain embodiments, those features may also be present in other embodiments whether described or not. Whilst endeavouring in the foregoing specification to draw attention to those features of the invention believed to be of particular importance it should be understood that the applicant claims protection in respect of any patentable feature or combination of features hereinbefore referred to and/or shown in the drawings whether or not particular emphasis has been placed thereon. l/we claim:

Claims

1. A rocket motor for a vertical take-off and landing (VTOL) aircraft, the rocket motor comprising:

a casing defining an elongate chamber for storing propellant, the elongate chamber having a length dimension, a width dimension and a depth dimension, wherein the length dimension is greater than the width dimension and greater than the depth dimension;

propellant, located in the elongate chamber, configured to enable the rocket motor to generate thrust in a direction that is substantially perpendicular to the length dimension of the elongate chamber by causing ejection of gas through a plurality of gas efflux apertures in the casing; and

at least one fluid injector for injecting a fluid fire suppressant into the elongate chamber.

2. The rocket motor of claim 1 , wherein the at least one fluid injector is for injecting a fluid fire suppressant into the elongate chamber, while thrust is being generated by the rocket motor, in order to increase the thrust.

3. The rocket motor of claim 1 or 2, wherein the at least one fluid injector is for injecting a fluid fire suppressant into the elongate chamber after thrust generation by the rocket motor has ceased.

4. The rocket motor of claim 1 , 2 or 3, wherein the at least one fluid injector is for injecting the fluid fire suppressant into a region of the elongate chamber that is located between the propellant and the plurality of gas efflux apertures.

5. The rocket motor of claim 4, wherein the casing further comprises a conduit, located in the region of the elongate chamber between the propellant and the plurality of gas efflux apertures in the casing, into which the fluid fire suppressant is injected.

6. The rocket motor of claim 5, wherein the conduit is an elongate conduit extending along the length dimension of the elongate chamber.

7. The rocket motor of claim 5 or 6, wherein the conduit comprises a plurality of fluid diffusing apertures arranged to release the fluid fire suppressant from the conduit within the region of the elongate chamber located between the propellant and the plurality of gas efflux apertures.

8. The rocket motor of claim 5, 6 or 7, further comprising an initiator, wherein the conduit is a pyrotechnic ignition conduit that enables ignition of the propellant, following ignition of the initiator.

9. The rocket motor of any of the preceding claims, wherein the fluid fire suppressant comprises one or more inert gases, carbon dioxide, water, water plus anti- freeze, halon, and/or one or more halocarbon-based agents.

10. The rocket motor of any of the preceding claims, wherein the at least one fluid injector comprises a control valve for controlling a flow rate of the fluid fire suppressant into the elongate chamber.

1 1. The rocket motor of any of the preceding claims, further comprising: one or more heaters arranged to heat the fluid fire suppressant.

12. The rocket motor of claim 11 , wherein the one or more heaters are configured to heat the fluid fire suppressant prior to injection of the fluid fire suppressant into the elongate chamber.

13. A VTOL aircraft comprising the rocket motor of any of the preceding claims.

14. An apparatus comprising the rocket motor of any of claims 1 to 12 and further comprising: control circuitry for controlling the at least one fluid injector.

15. The apparatus of claim 14, wherein the control circuitry is configured to control the at least one fluid injector to inject the fluid fire suppressant into the elongate chamber, while thrust is being generated by the rocket motor.

16. The apparatus of claim 14 or 15, wherein the control circuitry is configured to continue to control the at least one fluid injector to inject the fluid fire suppressant into the elongate chamber, after thrust generation by the rocket motor has ceased.

17. The apparatus of claim 14, 15 or 16, further comprising sensor circuitry for sensing attitude, altitude, drift and/or descent rate of a VTOL aircraft, wherein the control circuitry is configured to control the at least one fluid injector based, at least in part, on one or more inputs from the sensor circuitry.

18. A VTOL aircraft comprising the apparatus of any of claims 14 to 17.

19. A method, comprising:

analysing one or more inputs quantifying an attitude, altitude, drift and/or descent rate of a VTOL aircraft; and

controlling a flow rate of fluid fire suppressant into an elongate chamber of a rocket motor based, at least in part, on the analysis of the inputs.

20. Computer program code that, when executed by at least one processor, causes the method of claim 19 to be performed.

21. An apparatus, comprising:

at least one processor; and

memory storing computer program code that, when executed by the at least one processor, causes the apparatus to:

analyse one or more inputs quantifying an attitude, altitude, drift and/or descent rate of a VTOL aircraft; and

control a flow rate of fluid fire suppressant into an elongate chamber of a rocket motor based, at least in part, on the analysis of the inputs.

22. A rocket motor for a vertical take-off and landing (VTOL) aircraft, the rocket motor comprising:

propellant, located in the elongate chamber, configured to enable the rocket motor to generate thrust in a direction that is substantially perpendicular to the length dimension of the elongate chamber by causing ejection of gas through a plurality of gas efflux apertures in the casing;

an elongate conduit, located in a region of the elongate chamber between the propellant and the plurality of gas efflux apertures in the casing, comprising a plurality of fluid diffusing apertures; and at least one fluid injector for injecting a fluid into the elongate conduit.

23. The rocket motor of claim 22, wherein the at least one fluid injector is for injecting a fluid into the elongate conduit, while thrust is being generated by the rocket motor, in order to increase the thrust.

24. The rocket motor of claim 22 or 23, wherein the fluid is a fire suppressant.

25. The rocket motor of claim 22 or 23, wherein the fluid is a liquid propellant.

26. A method, comprising:

controlling a flow rate of fluid into an elongate conduit extending along an elongate chamber of a rocket motor based, at least in part, on the analysis of the inputs.

27. The method of claim 26, wherein the elongate conduit is located between propellant in the elongate chamber of the rocket motor and a plurality of gas efflux apertures in a casing of the rocket motor.

28. Computer program code that, when executed by at least one processor, causes the method of claim 26 or 27 to be performed.

29. An apparatus, comprising:

at least one processor; and

control a flow rate of fluid into an elongate conduit extending along an elongate chamber of a rocket motor based, at least in part, on the analysis of the inputs.

30. A rocket motor for a vertical take-off and landing (VTOL) aircraft, the rocket motor comprising: a casing defining a chamber for storing propellant, the propellant being for generating thrust by causing ejection of gas through a plurality of gas efflux apertures in the casing; and