GB2518211A

GB2518211A - Evaporative wick/membrane rocket motor

Info

Publication number: GB2518211A
Application number: GB1316343.1A
Authority: GB
Inventors: Carolyn Billie Knight; Geoffrey Daly
Original assignee: Individual
Current assignee: Individual
Priority date: 2013-09-13
Filing date: 2013-09-13
Publication date: 2015-03-18
Anticipated expiration: 2033-09-13
Also published as: GB2518211B; GB201316343D0

Abstract

A bi-propellant rocket motor 15 for combusting fuel 16 and oxidiser 17 to produce thrust. The rocket motor has a tubular casing 18, a combustion chamber 3 disposed radially inwards of the casing 18, and a means for supplying fuel 404 to the combustion chamber. Also provided is a means for supplying oxidiser 402 to the combustion chamber. The combustion chamber includes a porous membrane 4 and in use, fuel is supplied to the combustion chamber via the membrane and combusts in oxidiser supplied to the combustion chamber. The wall of the combustion chamber may comprise the porous membrane. The tubular casing and the porous membrane may form a void 5 for receiving fuel supplied to the combustion chamber. A nozzle 1 may be connected to the combustion chamber. The means for supplying fuel to the combustion chamber may be configured such that a portion of the fuel supplied to the combustion chamber may remain unburnt and passes along an inner surface of the membrane and nozzle to reduce ablative erosion of the nozzle.

Description

Evaporative Wick/Membrane Rocket Motor This invention relates generally to rocket motors, and more especially to bi-propellant motors.

Rocket motors have taken many forms. Each form has limitations and advantages for efficient propulsion and use. Solid propellant rocket motors can be expensive and dangerous to manufacture and lack controllability. They cannot be throttled or shut down.

Bi-propellant liquid rocket motors of the type shown in Figure 1 are controllable and can be shut down, but are extremely complex and expensive to manufacture and operate.

They employ a combustion chamber 3 and a nozzle 1 which require complex cooling systems and pumps 23,24. Oxidiser is pumped 23 from an external tank 17 and mixed with fuel, also pumped 24 from an external tank 16. While very simple in concept, in practice liquid bi-propellant rocket motors are extremely complex and expensive to manufacture and employ.

Hybrid rocket motors of the type shown in Figure 2 using liquid oxidisers 17 and solid fuel grains 25 are less expensive and are controllable by an oxidiser valve? but have not been proven to be capable of performing well on scales that may be useful for launch propulsion.

Mono-propellant motors of the type shown in Figure 3 use a liquid propellant 17 that can be made to react energetically when passed through a catalyst 26 to provide thrust.

Mono-propellant motors are relatively low powered and are used mainly in orbital vehicles for thrust vectoring. They are not commonly used for launch propulsion.

There is a need for a less expensive and fully controllable rocket motor that has the duration of burn and level of performance that is useful for launch propulsion of a meaningful payload.

According to a first aspect of the present invention, there is provided a bi-propellant rocket motor for combusting fuel and oxidiser to produce thrust. The rocket motor has a tubular casing, a combustion chamber disposed radially inwards of the casing, means for supplying fuel to the combustion chamber and means for supplying oxidiser to the combustion chamber. The combustion chamber includes a porous membrane and in use, fuel is supplied to the combustion chamber via the membrane and combusts in oxidiser supplied to the combustion chamber. The use of a porous membrane allows this rocket motor to produce performance equivalent to the bi-propellant motor of the type shown in Figure 2 but with lower cost and complexity.

Preferably, a wall of said combustion chamber comprises said porous membrane.

Preferably, the tubular casing and the porous membrane form a void for receiving fuel supplied to the combustion chamber.

Preferably, a nozzle connected to the combustion chamber; Preferably, the means for supplying fuel to the combustion chamber is configured so that a portion of the fuel supplied to said combustion chamber remains unburned and passes along an inner surface of said membrane and said nozzle thereby reducing ablative erosion of said nozzle. The film-layer cooling of the nozzle using fuel prevents the nozzle from vaporising.

Preferably, the combustion chamber comprises an output end assembly connected to said tubular casing and an input end assembly connected to said tubular casing.

Preferably, the means for supplying fuel comprises a fuel tank first.

Preferably, the rocket motor includes a conduit connected between said void and a fuel tank wherein, in use, fuel passes from said void to said fuel tank thereby cooling the motor.

Preferably, the porous membrane is comprised of a metal or an alloy or is comprised of a ceramic material.

The present invention also provides a combustion chamber for a rocket motor of the type described above.

The evaporative wick/membrane rocket motor of the present invention will reduce the expense of producing and employing a powerful and reliable rocket motor that can be fully controlled and shut down.

This motor will produce similar performance to the more complex and expensive liquid bi-propellant rocket motors. Unlike solid propellant rocket motors, it does not require the handling of explosive materials and it can be throttled and shut down. Its performance greatly exceeds that of the presently available hybrid rocket motors. This rocket motor represents an opportunity for a less expensive and more reliable means of accelerating payloads, either for space-launch propulsion or for other ballistic applications.

The use of the membrane allows a simplified bi-propellant motor, which currently use multiple injectors, and in which much technology and expense goes into effective atomisation and mixing. Furthermore, they get very hot very quickly and expensive cooling jackets and other means are employed, which are usually cooled by circulation of liquid oxygen or other cryogenic liquid. These cooling systems cool the combustion chamber and the nozzle and are complex and expensive. The geometry is entirely different and the combustion chambers themselves are, again very expensive to produce and to perfect.

For a more complete explanation of the present invention and the technical advantages thereof, reference is now made to the following description and the accompanying drawings in which: Figure 1 shows a schematic representation of a liquid bi-propellant rocket motor; Figure 2 shows a schematic representation of a hybrid rocket motor; Figure 3 shows a schematic representation of mono-propellant rocket motor; Figure 4 shows a schematic representation of the porous membrane rocket motor of the present invention; Figure 5 shows a schematic representation of the input end-cap assembly; Figure 6 shows a schematic representation of the output end-cap assembly; Figure 7 shows a schematic representation of the graphite nozzle; Figure 8 shows a schematic representation of the porous membrane assembly; Figure 9 shows a schematic representation of the ablative seal; and Figure 10 shows a schematic representation of the rocket containment assembly.

The present invention makes us of a porous membrane use to introduce fuel to the combustion chamber.

Embodiments of the present invention and their technical advantages may be better understood by referring to Figure 4, which shows a rocket motor 15 comprised of a tubular casing 18 and having a porous wick/membrane 4 enclosing a combustion chamber 3 having a nozzle 1. Tubular casing 18 may be comprised of aluminium, titanium or composite materials to minimise weight. Porous membrane 4 is cylindrical and is contained inside motor casing 18 and is concentric therewith. Porous membrane 4 may be comprised of a variety of materials including porous metals! porous ceramics or other porous materials tolerant of high temperatures.

Combustion chamber 3 is connected via conduit 402 to an external tank 12, which tank 12 is adapted to hold a liquid oxidiser 17. An igniter 6 capable, in use, of igniting the liquid oxidiser is positioned at an end of combustion chamber 3 that is distal to nozzle 1.

A void 5 between membrane 4 and casing 18 is connected to a fuel tank 9 via conduit 404, which tank 9 is adapted to hold a liquid fuel 16. An external pressurising tank 11 is connected to fuel tank 9 via conduit 406 and is adapted to be able to drive fuel from fuel tank 9 into void 5 and across porous membrane 4.

In use, porous membrane 4 permits a controlled flow of fuel into combustion chamber 3. Opening a valve 7 to introduce the liquid oxidiser into combustion chamber 3 where it is ignited by pyrotechnic or other type of igniter 6 starts combustion of the fuel. As the rocket burns, fuel 16 passes across membrane 4 into combustion chamber 3 to sustain combustion. The pressurised evaporation of fuel passing through porous membrane 4 allows for a controllable, well-atomised mist that will readily mix with the oxidiser 17 and allow for efficient combustion.

Fuel contained in void 5 between porous membrane 4 and outer casing 18 may be recirculated by means of a pump 6 via conduit 408 and external tank 9 thereby providing means of cooling the motor during combustion. A small quantity of unoxidised fuel, shown by arrows 49, will pass along an inner surface of porous membrane 4 and along an inner surface of nozzle 1 reducing ablative erosion of the nozzle 1 and increasing the possible burn-duration of the motor.

Flow rate of fuel across porous membrane 4 is controlled, for example by digital controller 22, which adjusts flow of fuel 16 into motor 15 by means of a valve 10.

Combustion pressure is constantly measured by a pressure transducer 20 which sends a signal to digital controller 22 which in turn operates pressure control valve 10 to balance fuel pressure for optimum oxidiser to fuel ratios. A power level of motor 15 is controlled by manipulation of an oxidiser valve 7, and motor 15 may be shut down by closing the same valve 7.

Motor 15 also comprises an input end-cap assembly 48 and an output end-cap assembly 2.

Referring to Figure 5, input end-cap assembly 48 is adapted to hold a starter injector 19, an oxidiser injector 46, and igniter 6. The oxidiser injector 46 is typically inserted at the centre of the assembly by means of a screw thread. Igniter 6 and the starter injector 47 are inserted in the same manner. Igniter 6 may be a black-powder pyrotechnic, initiated by an electric current, a constant-spark igniter or a gaseous oxdiser/spark igniter. All of these methods of ignition are well understood Referring to Figure 6, output end-cap assembly 2 comprises nozzle 1, which is typically comprised of a graphite nozzle insert 41, a nozzle holder 40, which typically comprises a metal or metal alloy, a nozzle-end-cap 39 and an insulating layer of fire cement 38. The graphite nozzle is retained by the layer of fire cement 38. inserted between graphite nozzle insert 41 and nozzle holder 40. Figure 7 shows detail of a typical graphite nozzle insert 41.

Referring now to Figure 7, porous metal membrane 4 assembly is sealed against casing 18 and end-cap assemblies 2,48 by means of a series of 0' rings 13. These function to contain liquid fuel within the void 5 between porous membrane 4 and outer casing 18. The fuel is typically liquid kerosene 17.

Referring to Figure 9, an ablative seal 14 sits between the nozzle assembly shown in Figure 6 and porous membrane 4. An outside diameter of ablative seal 14 is substantially the same as an internal diameter of casing 18. Ablative seal 14 typically comprises high-density polyethylene and functions to insulate porous membrane 4 from the nozzle assembly shown in Figure 6. Ablative seal 14 also functions to prevent flame propagation to outer casing 18. Effective sealing is accomplished by the use of a high-temperature resistant silicone sealant of industry standard equivalent.

Referring to Figure 10, which shows how end-cap assemblies 2,48 are held on casing by a tight interference fit and further held in place with an external arrangement of end-plates 43 and tie-rods 44 between end-plates 43.

Materials used in the construction of rocket motor 15 are chosen to be of sufficient strength to effectively contain an internal pressure of typically 400 lbs per square inch with an additional safety margin of 75% above expected load. All valves, controls and supply tanks are constructed according to industry practice, or are readily purchased on the open market. The porous membrane may comprise metals, alloys or ceramic materials. Metals and alloys include stainless steel, Inconel, Monel, bronze, or titanium.

In use, supply tanks 9,12,11 are filled and pressurised and fitted to motor 15 and all control valves 7,10,21 closed. Igniter 6 is initiated, and after a period of time, typically two seconds, oxidiser control valve 21 to starter injector 19 is opened and motor 15 will commence to burn. After a further period ot time, again typically two seconds, oxidiser control valve 7 to the main injector 46 is opened and rocket motor 15 will commence to burn at full power.

Claims

Claims 1. A bi-propellant rocket motor for combusting fuel and oxidiser to produce thrust, said rocket motor comprising: a tubular casing; a combustion chamber disposed radially inwards of said casing, said combustion chamber comprising a porous membrane; means for supplying fuel to said combustion chamber via said porous membrane; and means for supplying oxidiser to said combustion chamber; wherein, in use, fuel supplied to said combustion chamber via said membrane combusts in oxidiser supplied to said combustion chamber.
2. A bi-propellant rocket motor according to claim 1, in which a wall of said combustion chamber comprises said porous membrane.
3. A bi-propellant rocket motor according to claim 1 or2, in which said tubular casing and said porous membrane form a void for receiving fuel supplied to said combustion chamber.
4. A bi-propellant rocket motor according to any preceding claim comprising a nozzle connected to said combustion chamber;
5. A bi-propellant rocket motor according to claim 4, in which said means for supplying fuel to said combustion chamber is configured so that a portion of said fuel supplied to said combustion chamber remains unburned and passes along an inner surface of said membrane and said nozzle thereby reducing ablative erosion of said nozzle
6. A bi-propellant rocket motor according to any preceding claim, in which said combustion chamber comprises an output end assembly connected to said tubular casing and an input end assembly connected to said tubular casing.
7. A bi-propellant rocket motor according to any preceding claim, in which said means for supplying fuel comprises a first conduit connected between said void and a fuel tank.
8. A bi-propellant rocket motor according to claim 7 comprising a second conduit connected between said void and a fuel tank wherein, in use, fuel passes from said void to said fuel tank thereby cooling the motor.
9. A bi-propellant rocket motor according to any preceding claim, in which said porous membrane is comprised of a metal or an alloy.
10. A bi-propellant rocket motor according to any of claims ito 9, in which said porous membrane is comprised of a ceramic material.i2. A bi-propellant rocket motor substantially as described herein with reference to Figures 4 to 10 of the drawings.13. A combustion chamber for a rocket motor of any of the preceding claims.